diff --git a/.gitlab-ci.yml b/.gitlab-ci.yml
index a08f136e45ce5b387aed3f256054e806aa1ed32a..f9bc5ec0eeeced485d0114b6f0fb68f92015c146 100644
--- a/.gitlab-ci.yml
+++ b/.gitlab-ci.yml
@@ -61,8 +61,7 @@ pages:
# Install pipenv to manage Python dependencies
- apk update && apk --no-cache add graphviz
- pip install pipenv
- - pipenv install # Install the dependencies from the Pipfile
- - pipenv run pip install mkdoxy # Install the mkdoxy plugin if needed
+ - pipenv install # Install the dependencies from the Pipfile and Pipfile.lock
- apt-get update
- apt-get install -y doxygen
- export DOXYGEN_BIN=/usr/bin/doxygen
@@ -83,4 +82,4 @@ pages:
- if: '$CI_COMMIT_BRANCH != $CI_DEFAULT_BRANCH' # Allow manual triggers on non-default branches
when: manual # Run only when triggered manually
- if: '$CI_PIPELINE_SOURCE == "trigger"' # Triggered by another pipeline
- when: on_success # Run if the source pipeline was successful
\ No newline at end of file
+ when: on_success # Run if the source pipeline was successful
diff --git a/Pipfile b/Pipfile
index e2b42ecd38bec14409ba92a6b282222c50c5ec69..259f4e085529237018fdf6e6ccb0d836dd4b1aca 100644
--- a/Pipfile
+++ b/Pipfile
@@ -7,6 +7,8 @@ name = "pypi"
mkdocs = "*"
mkdocs-material = "*"
mkdocs-glightbox = "*"
+mkdocs-site-urls = "*"
+mkdoxy = "*"
[dev-packages]
mkdocs = "*"
diff --git a/Pipfile.lock b/Pipfile.lock
index 146ff6db1761d011930198a429478ffbd475a172..c0be02830948f3765a09f8d1ae3c01331d41e206 100644
--- a/Pipfile.lock
+++ b/Pipfile.lock
@@ -1,7 +1,7 @@
{
"_meta": {
"hash": {
- "sha256": "5fea7d31e7a99ffb6bc6be5d229a52f08658f807b7a75251742107aff366537a"
+ "sha256": "ba21ff566edd33d7f2eb9f27d67fd37e2cdec33b04fb5680c464356679bc7bd0"
},
"pipfile-spec": 6,
"requires": {
@@ -18,123 +18,125 @@
"default": {
"babel": {
"hashes": [
- "sha256:33e0952d7dd6374af8dbf6768cc4ddf3ccfefc244f9986d4074704f2fbd18900",
- "sha256:7077a4984b02b6727ac10f1f7294484f737443d7e2e66c5e4380e41a3ae0b4ed"
+ "sha256:0c54cffb19f690cdcc52a3b50bcbf71e07a808d1c80d549f2459b9d2cf0afb9d",
+ "sha256:4d0b53093fdfb4b21c92b5213dba5a1b23885afa8383709427046b21c366e5f2"
],
- "markers": "python_version >= '3.7'",
- "version": "==2.13.1"
+ "markers": "python_version >= '3.8'",
+ "version": "==2.17.0"
},
"certifi": {
"hashes": [
- "sha256:539cc1d13202e33ca466e88b2807e29f4c13049d6d87031a3c110744495cb082",
- "sha256:92d6037539857d8206b8f6ae472e8b77db8058fec5937a1ef3f54304089edbb9"
+ "sha256:3d5da6925056f6f18f119200434a4780a94263f10d1c21d032a6f6b2baa20651",
+ "sha256:ca78db4565a652026a4db2bcdf68f2fb589ea80d0be70e03929ed730746b84fe"
],
"markers": "python_version >= '3.6'",
- "version": "==2023.7.22"
+ "version": "==2025.1.31"
},
"charset-normalizer": {
"hashes": [
- "sha256:06cf46bdff72f58645434d467bf5228080801298fbba19fe268a01b4534467f5",
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+ ],
+ "markers": "python_version >= '3.7'",
+ "version": "==3.4.1"
},
"click": {
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- "sha256:ca9853ad459e787e2192211578cc907e7594e294c7ccc834310722b41b9ca6de"
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],
"markers": "python_version >= '3.7'",
- "version": "==8.1.7"
+ "version": "==8.1.8"
},
"colorama": {
"hashes": [
@@ -153,93 +155,94 @@
},
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],
- "markers": "python_version >= '3.5'",
- "version": "==3.4"
+ "markers": "python_version >= '3.6'",
+ "version": "==3.10"
},
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+ "version": "==3.1.5"
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+ "sha256:eaa0a10b7f72326f1372a713e73c3f739b524b3af41feb43e4921cb529f5929a",
+ "sha256:eb7972a85c54febfb25b5c4b4f3af4dcc731994c7da0d8a0b4a6eb0640e1d178",
+ "sha256:ee55d3edf80167e48ea11a923c7386f4669df67d7994554387f84e7d8b0a2bf0",
+ "sha256:f3818cb119498c0678015754eba762e0d61e5b52d34c8b13d770f0719f7b1d79",
+ "sha256:f8b3d067f2e40fe93e1ccdd6b2e1d16c43140e76f02fb1319a05cf2b79d99430",
+ "sha256:fcabf5ff6eea076f859677f5f0b6b5c1a51e70a376b0579e0eadef8db48c6b50"
],
- "markers": "python_version >= '3.7'",
- "version": "==2.1.3"
+ "markers": "python_version >= '3.9'",
+ "version": "==3.0.2"
},
"mergedeep": {
"hashes": [
@@ -251,147 +254,177 @@
},
"mkdocs": {
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- "sha256:3b3a78e736b31158d64dbb2f8ba29bd46a379d0c6e324c2246c3bc3d2189cfc1",
- "sha256:eb7c99214dcb945313ba30426c2451b735992c73c2e10838f76d09e39ff4d0e2"
+ "sha256:7b432f01d928c084353ab39c57282f29f92136665bdd6abf7c1ec8d822ef86f2",
+ "sha256:db91759624d1647f3f34aa0c3f327dd2601beae39a366d6e064c03468d35c20e"
],
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- "version": "==1.5.3"
+ "markers": "python_version >= '3.8'",
+ "version": "==1.6.1"
+ },
+ "mkdocs-get-deps": {
+ "hashes": [
+ "sha256:162b3d129c7fad9b19abfdcb9c1458a651628e4b1dea628ac68790fb3061c60c",
+ "sha256:2bf11d0b133e77a0dd036abeeb06dec8775e46efa526dc70667d8863eefc6134"
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+ "markers": "python_version >= '3.8'",
+ "version": "==0.2.0"
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+ "sha256:392b34207bf95991071a16d5f8916d1d2f2cd5d5bb59ae2997485ccd778c70d9",
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],
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- "version": "==0.3.4"
+ "version": "==0.4.0"
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+ "sha256:71d90dbd63b393ad11a4d90151dfe3dcbfcd802c0f29ce80bebd9bbac6abc753",
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- "version": "==9.4.6"
+ "version": "==9.6.2"
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- "version": "==1.3"
+ "version": "==1.3.1"
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- "version": "==23.2"
+ "version": "==0.2.0"
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+ "sha256:d584d9ec91ad65861cc08d42e834324ef890a082e591037abe114850ff7bbc3e",
+ "sha256:d7fded462629cfa4b685c5416b949ebad6cec74af5e2d42905d41e257e0869f5",
+ "sha256:d84a1718ee396f54f3a086ea0a66d8e552b2ab2017ef8b420e92edbc841c352d",
+ "sha256:d8e03406cac8513435335dbab54c0d385e4a49e4945d2909a581c83647ca0290",
+ "sha256:e10ce637b18caea04431ce14fabcf5c64a1c61ec9c56b071a4b7ca131ca52d44",
+ "sha256:ec031d5d2feb36d1d1a24380e4db6d43695f3748343d99434e6f5f9156aaa2ed",
+ "sha256:ef6107725bd54b262d6dedcc2af448a266975032bc85ef0172c5f059da6325b4",
+ "sha256:efdca5630322a10774e8e98e1af481aad470dd62c3170801852d752aa7a783ba",
+ "sha256:f753120cb8181e736c57ef7636e83f31b9c0d1722c516f7e86cf15b7aa57ff12",
+ "sha256:ff3824dc5261f50c9b0dfb3be22b4567a6f938ccce4587b38952d85fd9e9afe4"
],
- "markers": "python_version >= '3.6'",
- "version": "==6.0.1"
+ "markers": "python_version >= '3.8'",
+ "version": "==6.0.2"
},
"pyyaml-env-tag": {
"hashes": [
@@ -403,155 +436,431 @@
},
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diff --git a/README.md b/README.md
index ee00062a82aa56bcd96d9af52004365b681fc032..e284dc76d334461f9941ff1ee81d72e01d2b380f 100644
--- a/README.md
+++ b/README.md
@@ -7,7 +7,7 @@ We are using `mkdocs` as the static site generator to generate the website from
You can preview the website after cloning it. First you need to install `mkdocs` and the used theme by:
```sh
-pip install mkdocs mkdocs-material mkdocs-glightbox mkdoxy
+pip install mkdocs mkdocs-material mkdocs-glightbox mkdoxy mkdocs-site-urls
```
Then you can change the directory to the repository and run
diff --git a/docs/assets/javascripts/mathjax.js b/docs/assets/javascripts/mathjax.js
new file mode 100644
index 0000000000000000000000000000000000000000..0679bf6cd918eea6a1b468830d62d664d9292541
--- /dev/null
+++ b/docs/assets/javascripts/mathjax.js
@@ -0,0 +1,19 @@
+window.MathJax = {
+ tex: {
+ inlineMath: [["$", "$"]],
+ displayMath: [["$$", "$$"]],
+ processEscapes: true,
+ processEnvironments: true
+ },
+ options: {
+ ignoreHtmlClass: ".*",
+ processHtmlClass: "mathjax-render"
+ }
+};
+
+document$.subscribe(() => {
+ MathJax.startup.output.clearCache()
+ MathJax.typesetClear()
+ MathJax.texReset()
+ MathJax.typesetPromise()
+})
\ No newline at end of file
diff --git a/docs/documentation/additional_software.md b/docs/documentation/additional_software.md
index f91054796ecf9e4fad8ff24c6a132e3a8093ea12..33fe250e8bc97b23be3783013bbb04cd9faef8e1 100644
--- a/docs/documentation/additional_software.md
+++ b/docs/documentation/additional_software.md
@@ -8,7 +8,7 @@ glightbox: false
---
## cpacsInterface
-{.overview-img align=left}
+{.overview-img align=left}
The **cpacsInterface** is an additional module of the UNICADO toolchain.
Its purpose is to transform the UNICADO aircraft XML file into CPACS format and vice Versa.
The module consists of two modules convertUNICADO2CPACS which is responsible for converting the UNICADO aircraft exchange file into a CPACS format document, and convertCPACS2UNICADO which does the exact opposite and convert the data of the CPACS file into UNICADO format file.
@@ -21,7 +21,7 @@ The module consists of two modules convertUNICADO2CPACS which is responsible for
---
## designEvaluator
-{.overview-img align=left}
+{.overview-img align=left}
The **deignEvaluator** can be used to perform all available analysis on a designed aircraft and create all available reports for it.
{.overview-item}
@@ -32,8 +32,8 @@ The **deignEvaluator** can be used to perform all available analysis on a design
---
## reportGenerator
-{.overview-img align=left}
-The program collects all :simple-latex: reports of the programs and compiles a total pdf-report.
+{.overview-img align=left}
+The program collects all :simple-latex: reports of the programs and compiles a total pdf-report.
{.overview-item}
|Module Version|Language|License|Documentation|
@@ -43,7 +43,7 @@ The program collects all :simple-latex: reports of the programs and compiles a t
---
## testFramework
-{.overview-img align=left}
+{.overview-img align=left}
The **testFramework** is the heart of the **UNICADO** test pipeline.
It can perform all required test at the different hierarchy levels.
It is mainly an automation tool written specifically for the **UNICADO** project.
diff --git a/docs/documentation/analysis/aerodynamic_analysis/aerodynamic_principles.md b/docs/documentation/analysis/aerodynamic_analysis/aerodynamic_principles.md
new file mode 100644
index 0000000000000000000000000000000000000000..b650a9f8e4be673ddcff66880ba1dc5cf77cd629
--- /dev/null
+++ b/docs/documentation/analysis/aerodynamic_analysis/aerodynamic_principles.md
@@ -0,0 +1,123 @@
+
+# Aerodynamic principles {#aerodynamicprinciples}
+
+All methods for calculationg the properties of an aircraft face a trade off between accuracy on one hand and complexity and computing effort on the other hand.
+
+A typical aircraft in UNICADO takes roughly 20 to 30 iterations to converge in the design loop.
+For each iteration, the full aerodynamic properties have to be calculated.
+To enable extensive design space exploration and optimization studies in a reasonable time frame, the whole design process in UNICADO should takes less then an hour.
+The aerodynamic analysis therefore should be finished in under a minute.
+As a consequence of this requirement, the preliminary aircraft design in general, including UNICADO, is limited to lower fidelity methods, ranging from semi-empirical formulas to analytical approaches.
+
+**aerodynamic_analysis** contains a set off different methods and will be expanded in future.
+
+
+## Methods
+
+Currently there are **methods** with differing levels of fidelity implemented. These methods are listed in the table below.
+
+| Aerodynamic value | Methods | Fidelity level | Application |
+|-------------------------------------------------|-------------------------------|-----------------------------------|-------------------------------------------|
+|Lift, induced drag and pitching moment | Lifting Line | analytical | Lifting surfaces in general |
+|Lift, induced drag and pitching moment with corrections for TAW | Lifting Line | analytical/semi-empirical | Wing and stabilizer for TAW |
+|Viscous drag | According to Raymer | semi-empirical | Lifting surfaces, fuselages and nacelles |
+|Wave drag | According to Mason | semi-empirical | Lifting surfaces |
+|High lift adaptions | According to Raymer and Howe | semi-empirical | TAW configuration |
+|Trim function | Linear interpolation | - | Trimming via all movable horizontal stabilizer |
+
+The aim is to extend the method set with new calculation methods of variing fidelites for conventional TAW and and conventional configurations like the BWB.
+
+## Strategies
+
+The methods shown above have certain limitations:
+- No method can provide all aerodynamic values needed
+- The methods are only valid for certain flight conditions and aircraft configurations
+- Most methods need other aerodynamic values as input for their calculation
+
+Because of these shortcommings, the engineer has to select a suitable set of methods for their aircraft and bundle them together into a **strategy**.
+Due to the complexitiy in the fields of aerodynamics, the individual methods cannot be pluged in and out of a strategy, rather the strategies are tailor made for a given case.
+For illustration, the default strategy for calculation of the polars for the TAW is explained in the next chapter.
+
+
+## Example strategy for tube and wing
+
+### Lifting Line
+Lifting Line is a method to calculate the lift distribution and the induced drag.
+For this purpose, the potential equations are used, i.e. the flow is simplified and assumed to be frictionless, rotationless and incompressible.
+The wing is reduced to its skeletal lines.
+This simplified geometry is divided into trapezoidal elementary wings, which are covered with free and bound vortices.
+A system of equations is constructed from the vortex system and the boundary conditions, the solution of which is used to calculate the lift distribution.
+For a more in-depth discussion, the dissertation by Horstmann (Horstmann 1987: Ein Mehrfach-Traglinienverfahren und seine Verwendung für Entwurf und Nachrechnung nichtplanarer Flügelanordnungen) is recommended.
+
+The following picture shows the lifitng surfaces of a typical TAW aircraft discretized into elementary wings according to the lifting line method:
+
+
+The Prandtl-Glauert transformation is applied to the polars from Lifting Line.
+Lift coefficients, induced drag and pitch moment coefficients are thus transformed to include the compressibility effects.
+The lift distribution calculated using lifting line agrees well with CFD results for both the conventional wing and the blended wing body.
+Since the concept of the induced drag is based on the lifting line theory, it cannot be validated by CFD methods, which are based on the Navier-Stokes-Equations.
+Several semi-empirical corrections are integrated into the lifting line methodology in UNICADO.
+Based on Roskam, induced drag is calculated for the fuselage and nacelles.
+The pitching moment is corrected for fuselage and nacelle influences based on Torenbeek (Torenbeek, E. - Advanced Aircraft Design, 2013, ISBN: 9781119969303).
+
+### Viscous drag according to Raymer
+The frictional drag/viscous drag/zero lift drag is calculated based on the method of Raymer (Raymer 1992: Aircraft Design: A Conceptual Approach, page 280 ff).
+Contrary to what the name suggests, the viscous drag also regards influences of the boundary layer, which makes validation by CFD calculations difficult.
+
+For this purpose, the aircraft is broken down into its individual components, whose drag is calculated from a form factor, interference factor, friction coefficient and the wetted area:
+
+$
+ C_{D0} = \frac{\sum(C_{fc}FF_{c}Q_{c}S_{wet,c})}{S_{ref}}+C_{Dmisc}+C_{DLP}
+$
+
+The form factors are calculated using semi-empirical formulas, the interference factors are derived from the recommendations in the text (page 284 f).
+The friction coefficient is derived from the flow around a flat plate and depends on the Reynolds number and the surface roughness.
+
+In addition to the drags for the individual components, a 'miscellaneous drag' is calculated.
+This includes resistance caused by gas entering and leaving the hull through leaks and resistance caused by antennas, protrusions and the like.
+In total, the viscous drag depends only on the geometry, Reynolds number and Mach number and is thus constant over an entire aircraft polar.
+A calibration method is built in which the viscous drag is calibrated using an exponential function based on the lift coefficient.
+Thus, the viscous drag slightly increases with increasing lift.
+
+### Wave drag according to Mason
+The wave drag is the pressure drag generated by the occurrence of a shock wave.
+A compression shock reduces the static pressure of the fluid, which results in the surface pressure at the trailing edge of the profile being weaker than at the leading edge.
+The wave drag therefore only occurs when a compression shock occurs.
+From flight data it could be deduced that with increasing Mach number the wave drag is only between 0 and 10 drag counts and increases slightly linearly up to a Mach divergence number, above which the wave drag increases exponentially.
+This behavior of the wave drag is approximated by a fourth degree polynomial.
+
+The following picture shows the drag creep in the flight test data of a DC-9-30, according to Gur, Full-COnfiguration Drag Estimation, 2010:
+
+
+To calculate the wave drag, the critical Mach number is required, which is calculated according to the Korn-Mason equation (Mason 1990: Analytic Models for Technology Integration in Arcraft Design).
+To calculate the critical Mach number, the wing sweep, the profile thickness ratio, the local lift coefficient and the "profile technology factor" are required.
+Two values ​​are given for the profile technology factor, 0.87 for conventional and 0.95 for transonic profiles.
+Since the local lift coefficient is included in the formula for the critical Mach number, the wing is divided into individual strips for the drag calculation.
+
+For each strip, the local critical Mach number and the local wave drag are calculated and then summed up.
+In Gur 2010: Full-Configuration Drag Estimation a simple, area-weighted summation over all wing strips is proposed.
+The wave drag is then calibrated like the viscous drag using an exponential function based on the lift coefficient.
+
+### High lift polars
+Analysis of the aircraft in high lift configurations, with extended leading and trailing edge high lift devices, poses difficulties, even in numerical or experimental setups.
+In the interest of saving computing time and ressources in the aerodynamic analysis, the only valid option is to rely on semi-empirical calculations.
+
+The high lift polars are calculated for the following cases:
+
+- Take Off
+
+- Take Off landing gear retracted
+
+- Climb
+
+- Approach
+
+- Approach with landing gear
+
+- Landing
+
+For this, the number, type, postions and areas of all leading and trailing edge devices are read in.
+The geometric parameters of the high lift devices are used to calculate a maximum lift coefficient and shifts of the drag and moment coefficients, based on a set of semi-empirical formulas.
+
+The following picture shows the shifts in lift and drag in the high lift polars for a typical short medium range passernger aircraft according to the method:
+
\ No newline at end of file
diff --git a/docs/documentation/analysis/aerodynamic_analysis/getting_started.md b/docs/documentation/analysis/aerodynamic_analysis/getting_started.md
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+# Getting started {#getting-started}
+This guide will show you the basic usage of **aerodynamic_analysis**. Following steps are necessary (if you are new to UNICADO check out the [settings and outputs](#settingsandoutputs) first!)
+
+## Step-by-step
+
+It is assumed that you have the `UNICADO Package` installed including the executables. In case you are a developer, you need to build the tool first (see [build instructions on UNICADO website](https://unicado.pages.rwth-aachen.de/unicado.gitlab.io/developer/build/cpp/)).
+
+1. Take an `aircraft_exchange_file` with a fully designed aircraft (fuselage, wing, empennage and nacelles already sized)
+2. Fill out the configuration file - change at least:
+ - in `control_settings`
+ - `aircraft_exchange_file_name` and `aircraft_exchange_file_directory` to your respective settings
+ - `console_output` at least to `mode_1`
+ - `plot_output` to false (or define `inkscape_path` and `gnuplot_path`)
+ - in `program_settings`
+ - `Trim` enable/disable and tune the trim calculations
+ - `FlightConditions`define your flight conditions with altitude and mach number
+ - The different methods, like `ViscDragRaymer` which are listed can be fine tuned, and customized
+ - Enable/disable and set individual calibration factors in the different methods and for the overall polars in `DragCorrection`
+3. Open terminal and run **aerodynamic_analysis**
+
+
+Following will happen:
+- you see output in the console window
+- csv- files containing the raw lift, drag and moment data for all calculations are created in the `aerodynamic_analysis` folder
+- results are saved via xml-file in the `/aircraft_exchange_file/aero_data` for later use in e.g. **mission_analysis**
diff --git a/docs/documentation/analysis/aerodynamic_analysis/img/Drag_creep.png b/docs/documentation/analysis/aerodynamic_analysis/img/Drag_creep.png
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diff --git a/docs/documentation/analysis/aerodynamic_analysis/index.md b/docs/documentation/analysis/aerodynamic_analysis/index.md
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+# Introduction {#mainpage}
+The tool aerodynamic_analysis is on of the core tools in UNICADO. The overall goal is to calculate the lift and drag for all flight phases ranging from take off to cruise and landing.
+The gool of the tool is to...
+- Enable aerodynamic analysis for conventional and unconventional aircraft configurations
+- calculate the lift to drag polars for all flight phases regarding the aircraft geometry and the altitude and flight speed
+
+
+The [getting started](getting_started.md) gives you a first insight in how to execute the tool and how it generally works. To understand how the aerodynamic analysis works in detail, the documentation is split into a [aerodynamic principles](aerodynamic_principles.md) and a [software architecture](software_architecture.md) section.
+
diff --git a/docs/documentation/analysis/aerodynamic_analysis/software_architecture.md b/docs/documentation/analysis/aerodynamic_analysis/software_architecture.md
new file mode 100644
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+# Software architecture {#softwarearchitecture}
+
+The software architecture is structured into various modules and packages, each handling specific tasks. Below is a description of the main components
+
+- strategies:
+ - **Strategies** define the procedure of calculating the polars by initializing the aircraft geometry, calling methods and copying and processing data.
+ - There are different strategies implemented, stored in the folders corresponding to the aircraft configuration (e.g., `taw`, `bwb`).
+ - Each Strategy has a corresponding `data.cpp` for reading and writing data into the `aircraft.xml` and a `config.cpp`file for reading from the `config.xml`.
+
+- methods:
+ - **Methods** are either derived from literature or rely on external calculation sofwares, data bases or surrogate models.
+ - Methods are structured in a general way, so that they can be accessed by all strategies ranging over different aircraft configurations.
+ - Methods are stored in the `methods` folder and need to be initialized, by **geometry input**, **flight conditions** and **input parameters** from the config file.
\ No newline at end of file
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+ transform="translate(164.63,163.34)"
+ stroke="none"
+ fill="#000000"
+ font-family="Helvetica"
+ font-size="11px"
+ text-anchor="start"
+ id="g144">
+ <text
+ id="text143"><tspan
+ font-family="Helvetica"
+ id="tspan143">Gust</tspan></text>
+ </g>
+ </g>
+ <g
+ fill="none"
+ color="#000000"
+ stroke="currentColor"
+ stroke-width="2"
+ stroke-linecap="butt"
+ stroke-linejoin="miter"
+ id="g146">
+ <path
+ stroke="#a1132e"
+ d="m 118.68,159.76 h 38.26"
+ id="path145" />
+ </g>
+ </g>
+ <g
+ id="g152">
+ <title
+ id="title147">Design Point</title>
+ <g
+ fill="none"
+ color="#ffffff"
+ stroke="#a1132e"
+ stroke-width="1"
+ stroke-linecap="butt"
+ stroke-linejoin="miter"
+ id="g148" />
+ <g
+ fill="none"
+ color="#000000"
+ stroke="currentColor"
+ stroke-width="1"
+ stroke-linecap="butt"
+ stroke-linejoin="miter"
+ id="g150">
+ <g
+ transform="translate(164.63,179.84)"
+ stroke="none"
+ fill="#000000"
+ font-family="Helvetica"
+ font-size="11px"
+ text-anchor="start"
+ id="g149">
+ <text
+ id="text148"><tspan
+ font-family="Helvetica"
+ id="tspan148">Design Point</tspan></text>
+ </g>
+ </g>
+ <g
+ fill="none"
+ color="#000000"
+ stroke="currentColor"
+ stroke-width="1"
+ stroke-linecap="butt"
+ stroke-linejoin="miter"
+ id="g151">
+ <use
+ xlink:href="#gpPt6"
+ transform="matrix(3.12,0,0,3.12,137.81,176.26)"
+ color="#0071bc"
+ id="use150" />
+ </g>
+ </g>
+ <g
+ fill="none"
+ color="#ffffff"
+ stroke="#0071bc"
+ stroke-width="2"
+ stroke-linecap="butt"
+ stroke-linejoin="miter"
+ id="g153" />
+ <g
+ fill="none"
+ color="#000000"
+ stroke="currentColor"
+ stroke-width="2"
+ stroke-linecap="butt"
+ stroke-linejoin="miter"
+ id="g154" />
+ <g
+ fill="none"
+ color="#000000"
+ stroke="#000000"
+ stroke-width="0.5"
+ stroke-linecap="butt"
+ stroke-linejoin="miter"
+ id="g155" />
+ <g
+ fill="none"
+ color="#000000"
+ stroke="currentColor"
+ stroke-width="0.5"
+ stroke-linecap="butt"
+ stroke-linejoin="miter"
+ id="g157">
+ <path
+ stroke="#000000"
+ d="m 103.99,45.01 v 489 h 620 v -489 z"
+ id="path155" />
+ <g
+ transform="rotate(-90,179.3,110.21)"
+ stroke="none"
+ fill="#000000"
+ font-family="Sans"
+ font-size="18px"
+ text-anchor="middle"
+ id="g156">
+ <text
+ id="text155"><tspan
+ font-family="Sans"
+ id="tspan155">Thrust to Weight [-]</tspan></text>
+ </g>
+ </g>
+ <g
+ fill="none"
+ color="#000000"
+ stroke="currentColor"
+ stroke-width="0.5"
+ stroke-linecap="butt"
+ stroke-linejoin="miter"
+ id="g160">
+ <g
+ transform="translate(413.99,577.36)"
+ stroke="none"
+ fill="#000000"
+ font-family="Sans"
+ font-size="18px"
+ text-anchor="middle"
+ id="g159">
+ <text
+ id="text159"><tspan
+ font-family="Sans"
+ id="tspan157">Wing Loading [N/m</tspan><tspan
+ font-family="Sans"
+ font-size="14.4px"
+ dy="-9"
+ id="tspan158">2</tspan><tspan
+ font-family="Sans"
+ font-size="18px"
+ dy="9"
+ id="tspan159">]</tspan></text>
+ </g>
+ </g>
+ <g
+ fill="none"
+ color="#000000"
+ stroke="currentColor"
+ stroke-width="0.5"
+ stroke-linecap="butt"
+ stroke-linejoin="miter"
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+ <g
+ fill="none"
+ color="#000000"
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+ id="g162" />
+ </g>
+</svg>
diff --git a/docs/documentation/analysis/constraint_analysis/index.md b/docs/documentation/analysis/constraint_analysis/index.md
new file mode 100644
index 0000000000000000000000000000000000000000..2bad38713602ef5d6bd9c3a5dbd4a89207075f79
--- /dev/null
+++ b/docs/documentation/analysis/constraint_analysis/index.md
@@ -0,0 +1,37 @@
+# Constraint Analysis {#mainpage}
+One of the essential aspects of aircraft design is to size the aircraft to meet the point performance requirement. This process involves the evaluation of the constraints within the Thrust to Weight and Wing Loading design space. As the tool is "sizing" the aircraft, it requires information about the aerodynamic performance of the aircraft and the change of weight throughout the mission. Therefore this tool gets executed at the end of each loop to correctly size the aircraft.
+
+# Module Configuration
+The module can be configured to meet specific user needs by selecting desired parameters within the program_settings section of the module config file.
+A summary of possible selections can be found below:
+
+- `method`: This defines the method of constraint analysis
+ - Energy_Based
+
+- `aero_method`: This defines the method of getting information about the aerodynamic characteristics of the aircraft
+ - Calculate_Polar: Calculates the aerodynamic performance based on simple quadratic fit
+ - Read_Polar: Reads the polar infomration from the output of aerodynamic_analysis
+
+- `Mach_TO`: The mach number at takeoff
+
+- `takeoff_climb_angle`: The takeoff climb angle for which the constraint shall be evaluated
+
+- `gust_speed`: The gust speed at takeoff for which the constraint shall be evaluated
+
+- `gust_load_factor`: The additional gust load factor for which the constraint shall be evaluated
+
+- `oswald_factor`: The Oswald factor for calculating the polar, effective only if the aero_method is selected to be Calculate_Polar
+
+- `climb_gradient_OEI`: The minimum climb rate required for which the constraint shall be evaluated, CS25 defines this parameter to be 2.4% for the second climb segment
+
+- `minimum_climb_rate`: The minimum climb rate required at the service ceiling, CS25 defines this parameter to be 100 ft/min which is equal to 0.508 m/s
+
+- `safety_factor`: The additional percentage increment that is added to the Thrust to Weight ratio
+
+# Module Output
+
+- `Updated Design Point`: An updated Thrust to Weight and Wing Loading pair
+
+- `Constraint Plot`: The constraint plot if plotting is enabled
+
+- `Design Points`: The list of design points throughout the loop, to check for convergence performance
\ No newline at end of file
diff --git a/docs/documentation/analysis/constraint_analysis/principles.md b/docs/documentation/analysis/constraint_analysis/principles.md
new file mode 100644
index 0000000000000000000000000000000000000000..db95518a9ea33ee27fe0fd47c4c70b100e65d211
--- /dev/null
+++ b/docs/documentation/analysis/constraint_analysis/principles.md
@@ -0,0 +1,268 @@
+# Constraint Analysis
+
+# Aerodynamic and Performance Equations for Constraint Analysis Module
+
+## Function
+
+Adjust the design point based on the point performance requirements.
+
+## Rationale
+
+The end-of-the-loop aircraft’s aerodynamic performance is different to that of the beginning-of-the-loop aircraft. Therefore, it is necessary to size the aircraft based on the most up to date data from the design loop. This way, the design is ensured to have the best fitting characteristics to the mission and performance requirements.
+
+## Logic
+
+Point performance requirements are evaluated as constraints within the $T_{SL}/W_{TO}$ – $W_{TO}/S_{Ref}$ design space. The point is selected to have the minimum $T_{SL}/W_{TO}$ possible value that lies in the feasible space.
+
+# Baseline Implementation
+The constraint analysis tool is established with an energy based approach which is coming from Jack D. Mattingly's Aircraft Engine Design book.
+
+## Derivation
+
+1\. Lift Equation (getCL):
+
+$ L=n\\cdot W=\\frac{1}{2}\\cdot\\rho\\cdot\\left(M\\cdot a\\right)^2\\cdot C_L\\cdot S $
+
+$ C_L=\\frac{n}{\\frac{1}{2}\\cdot\\rho\\cdot\\left(M\\cdot a\\right)^2}\\cdot\\left(W/S\\right) $
+
+2\. Drag Equation:
+
+$ D=\\frac{1}{2}\\cdot\\rho\\cdot\\left(M\\cdot a\\right)^2\\cdot C_D\\cdot S $
+
+where:
+
+$ C_D=f\\left(C_L\\right) $
+
+3\. Thrust-Difference Equation:
+
+$ T-D=W\\cdot\\frac{d}{dt}\\left(h+\\frac{\\left(M\\cdot a\\right)^2}{2\\cdot g}\\right) $
+
+4\. Weight and Thrust Relationships:
+
+$ W=\\beta\\cdot W_{TO} $
+
+$ T=\\alpha\\cdot T_{SL} $
+
+where:
+
+$ \\alpha=f\\left(M,h\\right) $
+
+5\. Substituted Thrust-Difference Equation:
+
+$ T-\\left(\\frac{1}{2}\\cdot\\rho\\cdot\\left(M\\cdot a\\right)^2\\cdot f\\left(C_L\\right)\\cdot S\\right)=W\\cdot\\frac{d}{dt}\\left(h+\\frac{\\left(M\\cdot a\\right)^2}{2\\cdot g}\\right) $
+
+## Final Substituted Equation
+
+$ \\frac{T_{SL}}{W_{TO}}=\\frac{\\beta\\cdot\\frac{d}{dt}\\left(h+\\frac{V^2}{2\\cdot g}\\right)+\\frac{1}{2}\\cdot\\rho\\cdot\\left(M\\cdot a\\right)^2\\cdot f\\left(C_L\\right)}{\\alpha\\cdot\\left(W_{TO}/S\\right)} $
+
+$ f\\left(C_L\\right)=C_D\\left(\\frac{\\beta\\cdot n}{\\frac{1}{2}\\cdot\\rho\\cdot\\left(M\\cdot a\\right)^2}\\cdot\\left(W_{TO}/S\\right)\\right) $
+
+## Parameter Meanings
+
+L: Lift (N)
+
+n: Load factor (dimensionless)
+
+W: Weight (N)
+
+Ï: Air density (kg/m^3)
+
+M: Mach number (dimensionless)
+
+a: Speed of sound (m/s)
+
+$C_L$: Lift coefficient (dimensionless)
+
+$C_D$: Drag coefficient (dimensionless)
+
+S: Reference area (m^2)
+
+T: Thrust (N)
+
+D: Drag (N)
+
+h: Altitude (m)
+
+g: Gravitational acceleration (9.81 m/s^2)
+
+β: Weight fraction (dimensionless)
+
+$W_{TO}$: Takeoff weight (N)
+
+α: Thrust fraction (dimensionless), α = f(M, h)
+
+$T_{SL}$: Sea-level thrust (N)
+
+$f(C_L)$: Functional relationship defining drag coefficient in terms of lift coefficient
+
+## Cases
+
+Equation 1 (constant_altitude_speed_cruise):
+
+$ \\frac{T_{SL}}{W_{TO}}=\\frac{\\beta}{\\alpha}\\cdot\\left(\\frac{0.5\\cdot\\rho\\cdot\\left(M\\cdot a\\right)^2}{\\beta}\\cdot\\frac{1.0}{W/S}\\right)\\cdot C_D $
+
+Equation 2 (constant_speed_climb):
+
+$ \\frac{T_{SL}}{W_{TO}}=\\frac{\\beta}{\\alpha}\\cdot\\left(\\frac{C_D}{\\frac{\\beta}{q\\cdot W/S}}+\\frac{1}{u}\\cdot\\frac{dh}{dt}\\right) $
+
+where:
+
+$ u=M\\cdot a $
+
+$ q=0.5\\cdot\\rho\\cdot u^2 $
+
+Equation 3 (constant_altitude_speed_turn):
+
+$ \\frac{T_{SL}}{W_{TO}}=\\frac{\\beta}{\\alpha}\\cdot\\left(K_1\\cdot n^2\\cdot\\frac{\\beta}{0.5\\cdot\\rho\\cdot\\left(M\\cdot a\\right)^2}\\cdot W/S+\\frac{C_{D0}}{\\frac{\\beta}{0.5\\cdot\\rho\\cdot\\left(M\\cdot a\\right)^2}\\cdot W/S}\\right) $
+
+Equation 4 (horizontal_acceleration):
+
+$ \\frac{T_{SL}}{W_{TO}}=\\frac{\\beta}{\\alpha}\\cdot\\frac{0.5\\cdot\\rho\\cdot\\left(M\\cdot a\\right)^2}{\\beta}\\cdot\\frac{1.0}{W/S}\\cdot C_D\\cdot\\frac{1}{g_0}\\cdot\\frac{dv}{dt} $
+
+Equation 5 (takeoff_ground_roll):
+
+$ \\frac{T_{SL}}{W_{TO}} = \\frac{\\beta^2}{\\alpha} \\cdot \\frac{k_{T0}^2}{s_G \\cdot \\rho \\cdot g_0 \\cdot C_{L_{max,TO}}} \\cdot W/S $
+
+Equation 6 (braking_roll):
+
+$ W/S=\\frac{s_G\\cdot\\rho\\cdot\\ g_0\\cdot\\mathrm{\\ }\\ \\left(C_{D_L}-\\mu_B\\cdot\\ C_{L_L}\\right)}{\\beta\\cdot\\ l\\ n\\left(1+\\frac{C_{D_L}-\\mu_B\\cdot\\ C_{L_L}}{\\frac{\\mu_B\\cdot\\ C_{L_L}}{k_{T0}^2}}\\right)} $
+
+Equation 7 (service_ceiling):
+
+$ \\frac{T_{SL}}{W_{TO}}=\\frac{\\beta}{\\alpha}\\cdot\\left(K_1\\cdot\\frac{\\beta}{0.5\\cdot\\rho\\cdot\\left(M\\cdot a\\right)^2}\\cdot W/S+K_2+\\frac{C_{D0}}{\\frac{\\beta}{0.5\\cdot\\rho\\cdot\\left(M\\cdot a\\right)^2}\\cdot W/S}+\\frac{1}{M\\cdot a}\\cdot\\mathrm{SEP}\\ \\right) $
+
+Equation 8 (takeoff_climb_angle):
+
+$ \\frac{T_{SL}}{W_{TO}}=
+\\frac{\\beta}{\\alpha} \\cdot
+\\left(
+K_1 \\cdot \\frac{C_{L_{max,TO}}}{k_{T0}^2}
++ K_2
++ \\frac{C_{D0}}{\\frac{C_{L_{max,TO}}}{k_{T0}^2}}
++ \\sin \\gamma
+\\right) $
+
+Equation 9 (gust):
+
+$ W/S=\\frac{C_{L_\\alpha}\\cdot\\rho\\cdot V_{TO_L}\\cdot w_g}{d_{n_G}\\cdot2\\cdot\\beta} $
+
+## Building the Cases for Constraint Analysis
+
+Constraint 1: One Engine Inoperative
+
+Inputs are:
+
+W_over_S_data,
+
+CD_vector (taken from the polar or calculated based on quadratic fit),
+
+weight_fraction_TO (taken from the mission analysis results),
+
+alpha_TO (taken from the engine library and modified to account for OEI),
+
+M_TO,
+
+altitude = 0.0,
+
+load_factor = 1.0,
+
+2.4 / 100.0 \* climb_speed;
+
+Climb speed is taken from the aircraft XML file. The OEI climb gradient is 2.4%.
+
+Output is the minimum Thrust to Weight ratio that the aircraft shall have as a vector.
+
+Constraint 2: Service Ceiling (SEP)
+
+Inputs are:
+
+W_over_S_data,
+
+CD_vector (taken from the polar or calculated based on quadratic fit),
+
+weight_fraction_segment (taken from the mission analysis results),
+
+alpha_segment (taken from the engine library),
+
+M_max (taken from the aircraft XML file),
+
+altitude_cruise,
+
+load_factor = 1.0,
+
+climb_gradient = 0.508;
+
+The climb rate is 100 ft/min which corresponds to 0.508 m per second.
+
+Output is the minimum Thrust to Weight ratio that the aircraft shall have as a vector.
+
+Constraint 3: Landing Field Length
+
+Inputs are:
+
+CD_max_L (taken from the polar or calculated based on quadratic fit),
+
+CL_max_L (C_LmaxLanding node of the aircraft XML),
+
+weight_fraction_landing (taken from the mission analysis results),
+
+alpha_landing (taken from the engine library),
+
+M_TO (taken from the aircraft XML file),
+
+altitude = 0.0,
+
+my_B (breaking_coefficient node of the aircraft XML),
+
+s_G_L (takeoff or landing field length)
+
+Output is the maximum Wing Loading value that the aircraft shall have.
+
+Constraint 4: Gust
+
+Inputs are:
+
+CL_alpha (the slope of the linear segment of the CL-AoA polar),
+
+altitude = 0.0,
+
+V_TO_L (taken from the aircraft XML file),
+
+w_g (taken from the gust_speed node of the module config),
+
+dn_G (taken from the gust_load_factor node of the module config),
+
+weight_fraction_TO (taken from the mission analysis results),
+
+## Updating the Design Point
+
+Steps:
+
+1. Find the dominant curve: Takes the max $T_{SL}/W_{TO}$ required at each $W_{TO}/S_{Ref}$, hence gets the constraining curve
+2. Read the “safety factor†which adds the desired increment to the minimum $T_{SL}/W_{TO}$ for additional safety.
+3. Find the minimum $T_{SL}/W_{TO}$ from the dominant curve, get the $W_{TO}/S_{Ref}$ value corresponding to this $T_{SL}/W_{TO}$
+ 1. Not an “optimizationâ€, just a sorting algorithm
+ 2. Does not interpolate between points
+4. Update the design point
+
+## Functional Flow
+
+1. Initialize the Aircraft XML, Polar XML, Engine, and the Config XML
+2. Initialize the wing loading vector
+3. For each case:
+ 1. Read the polar that corresponds to the desired Mach number and the flight configuration
+ 2. Get the CL based on the Lift Equation
+ 3. Get the CD that corresponds to the CL by either:
+ 1. Getting the CD directly from the polar or,
+ 2. Calculating the CD based on quadratic fit
+ 4. Evaluate the constraint
+4. Assemble the constraints
+5. Find the dominant curve
+6. Evaluate the feasible area
+7. Find the minimum $T_{SL}/W_{TO}$ from the dominant curve and get the W/S that corresponds to this $T_{SL}/W_{TO}$
+8. Update the design point
+9. Plot the results
+10. Save the plots
+
+## Example Output
+
\ No newline at end of file
diff --git a/docs/documentation/analysis/cost_estimation/getting_started.md b/docs/documentation/analysis/cost_estimation/getting_started.md
index ff94edb83dc1c5aac7ce87e0d28da4ac32b1843a..7e1c20500c264e656b72a694e2c7e34eefa149d2 100644
--- a/docs/documentation/analysis/cost_estimation/getting_started.md
+++ b/docs/documentation/analysis/cost_estimation/getting_started.md
@@ -1,5 +1,6 @@
# Getting started
This section will guide you through the necessary steps to get the _cost\_estimation_ module up and running. It contains information on tool requirements and design parameters.
+
- [Aircraft exchange file](#aircraft-exchange-file) - Get information on necessary parameters from the _acXML_.
- [Module configuration file](#module-configuration-file) - Dive into cost estimation specific parameters.
- [Additional requirements](#additional-requirements) - Is anything else necessary to get the module running?
@@ -9,6 +10,7 @@ This section will guide you through the necessary steps to get the _cost\_estima
It is assumed that you have the `UNICADO package` installed including the executables and UNICADO libraries.
Generally, we use two files to set or configure modules in UNICADO:
+
- The aircraft exchange file (or _acXML_) includes
- data related inputs (e.g., aircraft configuration, transport task) and
- data related outputs (e.g., annual direct operating costs).
@@ -22,6 +24,7 @@ In the following sections you will find more information on how to configure the
Since the _cost\_estimation_ module is an assessment tool, it is assumed that a converged aircraft design and therefore all the necessary data are already available.
The following information is needed from the _acXML_:
+
1. Design specification
- Configuration information: Configuration type
- Transport task: Passenger definition, passenger class definition, and cargo definition
@@ -71,16 +74,13 @@ Program Settings
| | | | | | - Price per operating empty mass
| | | | | | - Rate insurance
| | | | | | - Rate interest
-| | | | | | - Residual value factor
-| | | | | - Crew
-| | | | | | - Salary variation
| | | | | - Flight cycles
-| | | | | | - Block time per flight
+| | | | | | - Block time supplement per flight
| | | | | | - Daily night curfew time
| | | | | | - Potential annual operation time
-| | | | | | - Annual lay days overhaul
-| | | | | | - Annual lay days reserve
-| | | | | | - Annual lay days maintenance
+| | | | | | - Annual lay hours overhaul
+| | | | | | - Annual lay hours reserve
+| | | | | | - Annual lay hours maintenance
| | | | | - Handling
| | | | | | - Fees handling
| | | | | - Landing
diff --git a/docs/documentation/analysis/cost_estimation/index.md b/docs/documentation/analysis/cost_estimation/index.md
index 7446bf117fbe7a4860861c4e3a898b394788cb54..3a2b1a6b891f5e64d556f9d5238d637d2e7a6ec5 100644
--- a/docs/documentation/analysis/cost_estimation/index.md
+++ b/docs/documentation/analysis/cost_estimation/index.md
@@ -1,5 +1,5 @@
# Introduction {#mainpage}
-Welcome to the _cost\_estimation_ module in UNICADO – where we take your aircraft operating costs from “hmm… probably a lot?†to laser-accurate precision! This tool is like a financial :crystal_ball for your aircraft, crunching numbers on fuel, maintenance, crew costs, and just about (almost) every other expense you can imagine. Think of it as your budgeting co-pilot, always ready to calculate so you can focus on the skies instead of spreadsheets. With _cost\_estimation_, you stay in control, keep the accountants happy, and land at your bottom line without any turbulence. So buckle up, and let’s start calculating!
+Welcome to the _cost\_estimation_ module in UNICADO – where we take your aircraft operating costs from “hmm… probably a lot?†to laser-accurate precision! This tool is like a financial :crystal_ball: for your aircraft, crunching numbers on fuel, maintenance, crew costs, and just about (almost) every other expense you can imagine. Think of it as your budgeting co-pilot, always ready to calculate so you can focus on the skies instead of spreadsheets. With _cost\_estimation_, you stay in control, keep the accountants happy, and land at your bottom line without any turbulence. So buckle up, and let’s start calculating!
## Summary of features
Here’s a quick rundown of what the tool currently does, along with a sneak peek at what's planned:
@@ -15,10 +15,12 @@ Blended-wing-body |... |... |under development
## A user's guide to cost calculation
The _cost\_estimation_ tool is your key to accurately calculating the operating costs of an aircraft. In this user documentation, you’ll find all the information you need to understand the tool, as well as the necessary inputs and configurations to run a cost analysis from the ground up.
The following sections will walk you through the cost estimation process in UNICADO:
+
- [Getting started](getting_started.md)
- [Run your first cost estimation](run_your_first_cost_estimation.md)
For a comprehensive understanding of the tool’s functionality, the documentation is structured into two distinct sections:
+
- A [method description](operating_cost_method.md) and
- a [software architecture](software_architecture.md)
section.
diff --git a/docs/documentation/analysis/cost_estimation/operating_cost_method.md b/docs/documentation/analysis/cost_estimation/operating_cost_method.md
index c3231270e1742732b946a42bfe62d21963336dde..9a8afa6f99cc846952f40779117ef9bba588eda7 100644
--- a/docs/documentation/analysis/cost_estimation/operating_cost_method.md
+++ b/docs/documentation/analysis/cost_estimation/operating_cost_method.md
@@ -1,88 +1,109 @@
# Calculation method
The total operating costs of an aircraft are split into direct operating costs (DOC) and indirect operating costs (IOC).
-$
+$$
TOC = DOC + IOC
-$
+$$
!!! note
- Unless explicitly stated, all values are in SI units and all costs in EUR.
+ Unless explicitly stated, all values are in SI units and all costs in EUR.
-## Direct operating costs (calculate_direct_operating_costs function)
+## Direct operating costs
The Direct Operating Costs (DOC) are directly influenced by the parameters and the aircraft's performance and are commonly used for aircraft evaluation. Therefore, a simplified method for DOC estimation, based on „From Aircraft Performance to Aircraft Assessment“ by J. Thorbeck <sup>[1]</sup>, is provided. The DOC are determined for one year and the entire depreciation period.
+
Two elements are required for the simplified DOC model: The route independent (fixed) costs $C_1$ and route dependent (variable) costs $C_2$:
-$
+$$
DOC = C_1 + C_2
-$
+$$
-### Route independent costs (calculate_route_independent_costs function)
+### Route independent costs
Route-independent costs include all cost components apart from the operation of the aircraft.
Hence, the route-independent costs are the sum of the capital costs and the crew costs:
-$
- C_1 = C_{CAP} + C_{crew}
-$
+$$
+ C_1 = C_{\text{capital}} + C_{\text{crew}}
+$$
+
Those are calculated both, for one year and for the depreciation period.
-#### Capital costs (calculate_capital_costs function)
+#### Capital costs
The capital costs can be assumed to be a linear function of the operating empty mass if the influence of the aircraft market is considered negligible:
-$
- C_{CAP} = P_{OE} \cdot m_{OE} \cdot (a+f_I)
-$
+$$
+ C_{\text{capital}} = P_{\text{OE}} \cdot m_{\text{OE}} \cdot (a + f_{\text{I}})
+$$
+
In which
-- $P_{OE}$ - price per kg operating empty mass
-- $m_{OE}$ - operating empty mass
+
+- $P_{\text{OE}}$ - price per kg operating empty mass
+- $m_{\text{OE}}$ - operating empty mass
- $a$ - annuity factor in percent
-- $f_I$ - insurance rate in percent
+- $f_{\text{I}}$ - insurance rate in percent
The annuity formula, which is based on a modified mortgage equation, addresses both yearly depreciation and interest:
-$
- a = f_{IR} \cdot \frac{1-f_{RV} \cdot \left(\frac{1}{1+f_{IR}}\right)^{t_{DEP}}}{1-\left(\frac{1}{1+f_{IR}}\right)^{t_{DEP}}}
-$
+$$
+ a = f_{\text{IR}} \cdot \frac{1 - f_{\text{RV}} \cdot \left( \frac{1}{1 + f_{\text{IR}}} \right)^{t_{\text{DEP}}}}{1 - \left( \frac{1}{1 + f_{\text{IR}}} \right)^{t_{\text{DEP}}}}
+$$
+
In which
-- $f_{IR}$ - interest rate in percent
-- $f_{RV}$ - residual value factor in percent
-- $t_{DEP}$ - depreciation period in years
+
+- $f_{\text{IR}}$ - interest rate in percent
+- $f_{\text{RV}}$ - residual value factor in percent
+- $t_{\text{DEP}}$ - depreciation period in years
The reason for the annuity method modification is to include the residual aircraft value at the end of the depreciation period into the capital costs, which is occasionally relevant. This assumes that an operator is purchasing an aircraft at a constant price per kilogram and spends the corresponding capital cost consistently per year throughout the depreciation period.
-#### Crew costs (calculate_crew_costs function)
+The residual value factor depends on the depreciation period and can be determined based on the following information:
+
+Depreciation period | Residual value factor |
+----------------------------|:---------------------:|
+up to 5 years | 0.7% |
+up to 10 years | 0.5% |
+up to 15 years | 0.3% |
+more than 15 years | 0.1% |
+
+#### Crew costs
This method is based on the lecture "J Flugzeugbewertung" by A. Bardenhagen <sup>[2]</sup>.
The annual crew costs are assumed to be the sum of the flight and cabin crew costs:
-$
- C_{crew} = C_{FC} + C_{CC}
-$
+$$
+ C_{\text{crew}} = C_{\text{FC}} + C_{\text{CC}}
+$$
both of which are of different levels. There are different approaches here, which must be adapted to the respective cost structure of the airline:
+
- Some airlines (mainly low-cost carriers) employ and pay pilots and flight attendants on a time basis (block hours).
- Other airlines hire their personnel permanently and must pay them irrespective of the time they are deployed.
-In the first case, the personnel costs belong to the variable in the second case to the fixed direct operating costs. Here, crew costs are assumed to be fixed (route independent) because an airline must provide enough crews to ensure flight operations over the entire service time and therefore are proportional to the payload. 50 passengers per flight attendant are assumed based on certification requirements. Crew costs are constant per year. To calculate the crew cost for several years, the expected salary increase should be considered by an escalation factor. Accordingly, past price levels can be extrapolated to the current level changed according to inflation, price, or salary increase.
+In the first case, the personnel costs belong to the variable in the second case to the fixed direct operating costs. Here, crew costs are assumed to be fixed (route independent) because an airline must provide enough crews to ensure flight operations over the entire service time and therefore are proportional to the payload. 50 passengers per flight attendant are assumed based on certification requirements.
+
+Crew costs are constant per year. To calculate the crew cost for several years, the expected salary increase should be considered by an escalation factor. Accordingly, past price levels can be extrapolated to the current level changed according to inflation, price, or salary increase.
Both cost shares are determined by the same variables:
-- The flight/cabin crew complement (the number of crews per aircraft, dependent on the stage length): $n_{FCC}$/$n_{CCC}$,
-- The number of flight/cabin crew members: $n_{FC}$/$n_{CC}$,
-- The annual salary of a flight/cabin crew member (dependent on the stage length): $S_{FC}$/$S_{CC}$, and
-- The escalation factor in percent: $f_{ESC}$.
+
+- The flight/cabin crew complement (the number of crews per aircraft, dependent on the stage length): $n_{\text{FCC}}$/$n_{\text{CCC}}$,
+- the number of flight/cabin crew members: $n_{\text{FC}}$/$n_{\text{CC}}$,
+- the annual salary of a flight/cabin crew member (dependent on the stage length): $S_{\text{FC}}$/$S_{\text{CC}}$, and
+- the escalation factor in percent: $f_{\text{ESC}}$.
<!-- NOTE: The values of these drivers depend on the stage length. Two modes are implemented. Mode 1 (salary_variation = False, default): To ensure that the values of the above-mentioned parameters are the same for the design mission and mission study, the stage length of the design mission is used to determine the values for the study mission as well. Mode 2 (salary_variation = True): The above-mentioned values are obtained for different stage lengths for the design mission and mission study. -->
That results in the following calculations:
-$
- C_{FC} = n_{FCC} \cdot n_{FC} \cdot S_{FC} \cdot f_{ESC}
-$
-$
- C_{CC} = n_{CCC} \cdot n_{CC} \cdot S_{CC} \cdot f_{ESC}
-$
+$$
+ C_{\text{FC}} = n_{\text{FCC}} \cdot n_{\text{FC}} \cdot S_{\text{FC}} \cdot f_{\text{ESC}}
+$$
+
+$$
+ C_{\text{CC}} = n_{\text{CCC}} \cdot n_{\text{CC}} \cdot S_{\text{CC}} \cdot f_{\text{ESC}}
+$$
The escalation factor
-$
- f_{ESC} = (1+r_{INF})^{y}
-$
+$$
+ f_{\text{ESC}} = (1 + r_{\text{INF}})^{y}
+$$
-incorporates the inflation rate ($r_{INF}$), which encompasses both price and salary adjustments, and the number of years elapsed between the calculation year and the base year for salaries ($y$).
+incorporates the inflation rate ($r_{\text{INF}}$), which encompasses both price and salary adjustments, and the number of years elapsed between the calculation year and the base year for salaries ($y$).
If the depreciation period is used as the time difference, resulting costs are related to the whole depreciation period, whereas a time difference of one year solely results in the costs for the base year.
The crew complements as well as the average annual salaries are dependent on the stage length:
+
- Regional: ranges less than 500 km
- Short haul: ranges between 500 km and 1000 km
- Medium haul: ranges between 1000 km and 4000 km
@@ -91,28 +112,30 @@ The crew complements as well as the average annual salaries are dependent on the
and can be taken from the following tables:
-Segment | Crew complement | $S_{FC}$ in EUR/y | $S_{CC}$ in EUR/y |
-----------------|:---------------:|:-----------------:|:-----------------:|
-Regional | 5 | 70 000 | 30 000 |
-Short haul | 5 | 120 000 | 30 000 |
-Medium haul | 5 | 160 000 | 30 000 |
-Long haul | 8 | 200 000 | 45 000 |
-Ultra-long haul | 8 | 200 000 | 45 000 |
+Segment | Crew complement | $S_{\text{FC}}$ in EUR/y | $S_{\text{CC}}$ in EUR/y |
+----------------|:---------------:|:------------------------:|:------------------------:|
+Regional | 5 | 70,000 | 30,000 |
+Short haul | 5 | 120,000 | 30,000 |
+Medium haul | 5 | 160,000 | 30,000 |
+Long haul | 8 | 200,000 | 45,000 |
+Ultra-long haul | 8 | 200,000 | 45,000 |
-### Route dependent costs (calculate_route_dependent_costs function)
+### Route dependent costs
Route dependent costs $C_2$ include all cost components that are directly attributable to flight operations. These include
-- fuel $C_F$,
-- fees (handling $C_H$, landing $C_L$, air traffic control (ATC) $C_{ATC}$), and
-- maintenance $C_{MRO}$.
+
+- fuel $C_\text{F}$,
+- fees (handling $C_\text{H}$, landing $C_\text{LDG}$, air traffic control (ATC) $C_{\text{ATC}}$), and
+- maintenance $C_{\text{MRO}}$.
Thus, the **annual** route dependent costs can be calculated by
-$
- C_2 = C_F + C_H + C_{LDG} + C_{ATC} + C_{MRO}
-$
+$$
+ C_2 = C_\text{F} + C_\text{H} + C_\text{LDG} + C_{\text{ATC}} + C_{\text{MRO}}
+$$
#### Flights per year
Knowing the number of annual flights is mandatory to calculate the above-mentioned cost shares.
A reliable approximation of the number of annual flights can be found using the following analytical basis:
+
- Potential flight hours per year: $365 \cdot 24 = 8760$
- Maintenance lay days per year (C-Check every 15 months for 4 days): $4 \cdot 12/15 = 3.2$
- Overhaul lay days per year (D-Check every 5 years for 4 weeks): $4 \cdot 7/5 = 5.6$
@@ -124,168 +147,182 @@ A reliable approximation of the number of annual flights can be found using the
- Yearly operation time in hours: $OT = 8760-2475-273.6 = 6011.4$
Knowing the time for one flight $FT$ and the block time supplement $BT$ (turn around time) per flight, the number of flight cycles $FC$ can be calculated:
-$
+$$
FC = \frac{OT}{(FT + BT)}
-$
+$$
It is assumed that one flight cycle consists of an outbound flight, a turnaround time and a return flight. Consequently, the number of annual flights is calculated as follows:
-$
- n_{flights} = 2 \cdot FC
-$
-
-#### Fuel costs (calculate_fuel_costs function)
-The fuel costs depend on the fuel price $P_F$, the trip fuel mass $m_{TF}$ (which can be obtained from the payload range diagram (PRD)), and the number of yearly flights $n_{flights}$:
-$
- C_F = P_{F} \cdot m_{TF}Â \cdot n_{flights}
-$
-
-#### Handling costs (calculate_handling_costs function)
-Handling charges $F_H$ include charges for loading and unloading, use of terminals and passenger boarding bridges, security checks, and ground energy supply.
-The annual handling fees are charged based on the payload mass $m_{PL}$ and the number of flights per year. The resulting handling costs are calculated as follows:
-$
- C_H = m_{PL} \cdot F_{H} \cdot n_{flights}
-$
-
-#### Landing costs (calcutale_landing_costs function)
-The annual landing fees $F_{LDG}$ are charged based on the maximum (certified) takeoff mass $m_{TO}$ and number of flights per year. The resulting landing costs are calculated as follows:
-$
- C_{LDG} = m_{TO} \cdot F_L \cdot n_{flights}
-$
-
-#### Air traffic control costs (calculate_air_traffic_control_costs function)
+$$
+ n_{\text{flights}} = 2 \cdot FC
+$$
+
+#### Fuel costs
+The fuel costs depend on the fuel price $P_\text{F}$, the trip fuel mass $m_{\text{TF}}$ (which can be obtained from the payload range diagram (PRD)), and the number of yearly flights $n_{\text{flights}}$:
+$$
+ C_\text{F} = P_{\text{F}} \cdot m_{\text{TF}}Â \cdot n_{\text{flights}}
+$$
+
+#### Handling costs
+Handling charges $F_\text{H}$ include charges for loading and unloading, use of terminals and passenger boarding bridges, security checks, and ground energy supply.
+The annual handling fees are charged based on the payload mass $m_{\text{PL}}$ and the number of flights per year. The resulting handling costs are calculated as follows:
+$$
+ C_\text{H} = m_{\text{PL}} \cdot F_{\text{H}} \cdot n_{\text{flights}}
+$$
+
+#### Landing costs
+The annual landing fees $F_{\text{LDG}}$ are charged based on the maximum (certified) takeoff mass $m_{\text{TO}}$ and number of flights per year. The resulting landing costs are calculated as follows:
+$$
+ C_{\text{LDG}} = m_{\text{TO}} \cdot F_\text{L} \cdot n_{\text{flights}}
+$$
+
+#### Air traffic control costs
The calculation of the ATC costs is based on the EUROCONTROL route charge formula <sup>[3]</sup>, more precisely the aircraft weight factor.
> "The weight factor (expressed to two decimals) is determined by dividing, by fifty (50), the certificated Maximum Take-Off Weight (MTOW) of the aircraft (in metric tonnes, to one decimal) and subsequently taking the square root of the result rounded to the second decimal [...]".
-The ATC price factor $f_{ATC}$ considers the fact that the price scenarios are varying strongly for each continent (or even region):
-- $f_{ATC} = 1.0$ for domestic europe
-- $f_{ATC} = 0.7$ for transatlantic flights
-- $f_{ATC} = 0.6$ for far east flights (only half of the landings at european airports)
+The ATC price factor $f_{\text{ATC}}$ considers the fact that the price scenarios are varying strongly for each continent (or even region):
+
+- $f_{\text{ATC}} = 1.0$ for domestic europe
+- $f_{\text{ATC}} = 0.7$ for transatlantic flights
+- $f_{\text{ATC}} = 0.6$ for far east flights (only half of the landings at european airports)
The ATC costs are calculated as follows:
-$
- C_{ATC} = R \cdot f_{ATC} \cdot \sqrt{\frac{m_{TO}[\text t]}{50}} \cdot n_{flights}
-$
+$$
+ C_{\text{ATC}} = R \cdot f_{\text{ATC}} \cdot \sqrt{\frac{m_{\text{TO}}[\text t]}{50}} \cdot n_{\text{flights}}
+$$
with
+
- $R$ - range in km
-- $m_{TO}$ - maximum takeoff mass (in tonnes)
+- $m_{\text{TO}}$ - maximum takeoff mass (in tonnes)
-#### Maintenance costs (calculate_maintenance_costs function)
+#### Maintenance costs
Maintenance costs are categorized into three components:
+
- Flight cycle dependent cost: This component primarily accounts for structural fatigue and overhaul burdens.
- Flight hour dependent cost: This component primarily reflects wear and the associated line maintenance work.
- Calendar time dependent cost: This component represents a constant share, such as the rectification of corrosion during overhaul.
In the following, only the maintenance costs per flight cycle are considered. Following the JADC method, an approximation for those costs is given by the sum of three parts:
-- Airframe material maintenance cost (repair and replacement): $C_{MRO,AF,MAT}$
-- Airframe personnel maintenance cost (inspection and repair): $C_{MRO,AF,PER}$
-- Engine total maintenance cost: $C_{MRO,ENG}$
+
+- Airframe material maintenance cost (repair and replacement): $C_{\text{MRO,AF,MAT}}$
+- Airframe personnel maintenance cost (inspection and repair): $C_{\text{MRO,AF,PER}}$
+- Engine total maintenance cost: $C_{\text{MRO,ENG}}$
In which
-$
- C_{MRO,AF,MAT} = m_{OE}[\text t] \cdot (0.2 \cdot t_{flight} + 13.7) + C_{MRO,AF,REP}
-$
-$
- C_{MRO,AF,PER} = f_{LR} \cdot (1+C_B) \cdot \left[ (0.655 + 0.01 \cdot m_{OE}[\text t]) \cdot t_{flight} + 0.254 + 0.01 \cdot m_{OE}[\text t] \right]
-$
-$
- C_{MRO,ENG} = n_{ENG} \cdot \left( 1.5 \cdot \frac{T_{0} [\text t]}{n_{ENG}} + 30.5 \cdot t_{flight} + 10.6 \cdot f_{MRO,ENG}\right)
-$
+$$
+ C_{\text{MRO,AF,MAT}} = m_{\text{OE}}[\text t] \cdot (0.2 \cdot t_{\text{flight}} + 13.7) + C_{\text{MRO,AF,REP}}
+$$
+
+$$
+ C_{\text{MRO,AF,PER}} = f_{\text{LR}} \cdot (1+C_\text{B}) \cdot \left[ (0.655 + 0.01 \cdot m_{\text{OE}}[\text t]) \cdot t_{\text{flight}} + 0.254 + 0.01 \cdot m_{\text{OE}}[\text t] \right]
+$$
+
+$$
+ C_{\text{MRO,ENG}} = n_{\text{ENG}} \cdot \left( 1.5 \cdot \frac{T_{0} [\text t]}{n_{\text{ENG}}} + 30.5 \cdot t_{\text{flight}} + 10.6 \cdot f_{\text{MRO,ENG}}\right)
+$$
with
-- $C_{MRO,AF,REP}$ - airframe repair cost per flight
-- $f_{LR}$ - labor rate in EUR/h
-- $C_B$ - cost burden
-- $n_{ENG}$ - number of engines
+
+- $C_{\text{MRO,AF,REP}}$ - airframe repair cost per flight
+- $f_{\text{LR}}$ - labor rate in EUR/h
+- $C_\text{B}$ - cost burden
+- $n_{\text{ENG}}$ - number of engines
- $T_{0}$ - sea level static thrust per engine
-- $f_{MRO,ENG}$ - engine maintenance factor
+- $f_{\text{MRO,ENG}}$ - engine maintenance factor
-The airframe repair cost per flight $C_{MRO,AF,REP}$ equal 57.5 for kerosene-powered aircraft. For hydrogen-powered aircraft, this value is multiplied by the operating empty mass factor $f_{OEM} = 1.1$ to account for an approx. 10% higher operating empty mass.
-The engine maintenance factor is considered $f_{ENG} = 1$ for kerosene-powered aircraft and $f_{ENG} = 0.7$ for hydrogen-powered aircraft.
+The airframe repair cost per flight $C_{\text{MRO,AF,REP}}$ equal `57.5` for kerosene-powered aircraft. For hydrogen-powered aircraft, this value is multiplied by the operating empty mass factor $f_{\text{OEM}} = 1.1$ to account for an approx. 10% higher operating empty mass.
+The engine maintenance factor is considered $f_{\text{ENG}} = 1$ for kerosene-powered aircraft and $f_{\text{ENG}} = 0.7$ for hydrogen-powered aircraft.
Thus, the annual maintenance costs result in
-$
- C_{MRO} = (C_{MRO,AF,MAT} + C_{MRO,AF,PER} + C_{MRO,ENG}) \cdot n_{flights}
-$
+$$
+ C_{\text{MRO}} = (C_{\text{MRO,AF,MAT}} + C_{\text{MRO,AF,PER}} + C_{\text{MRO,ENG}}) \cdot n_{\text{flights}}
+$$
## Related direct operating costs
Absolute DOC are generally unsuitable as an assessment measure because aircraft size and technology strongly influence this figure. They are therefore expressed in differently related quantities, depending on the purpose of the evaluation:
+
- DOC/Range (Flight Kilometer): Flight Kilometer Costs (FKC)
- DOC/Seat Kilometer Offered (SKO): Seat Kilometer Costs (SKC)
- DOC/Seat Kilometer Offered Corrected: Corrected SKC to take account of any freight revenue
- DOC/Ton Kilometers Offered (TKO): Ton Kilometer Costs (TKC)
- DOC/Revenue Passenger Kilometer (RPK): Revenue Seat Kilometer Costs (RSKC)
+
These are described below.
-### Flight kilometer costs (calculate_flight_kilometer_costs function)
+### Flight kilometer costs
The flight kilometer costs are very flexible and suitable for an extended consideration of changed route structures. This parameter allows the range potential of the aircraft to be assessed:
-$
+$$
FKC = \frac{DOC}{R}.
-$
+$$
+
+### Seat kilometer costs
+The seat kilometer offered (SKO) (or available) is a measure of an aircraft's passenger carrying capacity or, in other words, its potential to generate revenue by providing available seats to passengers. They are calculated by multiplying the number of seats available $n_{\text{seats}}$ by the range:
+$$
+ SKO = n_{\text{seats}} \cdot R.
+$$
-### Seat kilometer costs (calculate_seat_kilometer_costs function)
-The seat kilometer offered (SKO) (or available) is a measure of an aircraft's passenger carrying capacity or, in other words, its potential to generate revenue by providing available seats to passengers. They are calculated by multiplying the number of seats available $n_{seats}$ by the range:
-$
- SKO = n_{seats} \cdot R.
-$
The seat kilometer costs allow the analysis of a change in seat capacity and thus the assessment of the passenger kilometer potential:
-$
+$$
SKC = \frac{DOC}{SKO}
-$
+$$
-### Corrected seat kilometer costs (calculate_corrected_seat_kilometer_costs function)
+### Corrected seat kilometer costs
!!! note
- The calculation of this cost share is not implemented at the moment and set to `0` instead.
+ The calculation of this cost share is not implemented at the moment and set to `0` instead.
A method of freight equivalent passenger seats is applied.
-Cargo revenue from residual cargo payload at maximum zero fuel mass ($m_{PL,max} - m_{PL}$) can be calculated using
-$
- I_{cargo} = I_{FR} \cdot (W_{PL,max} - W_{PAX})
-$
+Cargo revenue from residual cargo payload at maximum zero fuel mass ($m_{\text{PL,max}} - m_{\text{PL}}$) can be calculated using
+$$
+ I_{\text{cargo}} = I_{\text{FR}} \cdot (W_{\text{PL,max}} - W_{\text{PAX}})
+$$
+
with
-- $I_{FR}$ - revenue per freight kilometer
-- $W_{PL,max}$ - maximum payload weight
-- $W_{PAX}$ - pax weight
+
+- $I_{\text{FR}}$ - revenue per freight kilometer
+- $W_{\text{PL,max}}$ - maximum payload weight
+- $W_{\text{PAX}}$ - pax weight
The equivalent seat revenue can be derived using the following formula:
-$
- n_{PAX,cargo} = \frac{I_{cargo}}{I_{PAX}}
-$
-with $I_{PAX}$ as revenue per seat and flight (see following table).
-
-Segment | $I_{PAX,multi-class}$ in EUR/SO | $I_{PAX,all-economy}$ in EUR/SO |
-----------------|:-------------------------------:|:-------------------------------:|
-Short haul | 400 | 250 |
-Medium haul | 450 | 300 |
-Long haul | 550 | 400 |
-Ultra long haul | 700 | 550 |
+$$
+ n_{\text{PAX,cargo}} = \frac{I_{\text{cargo}}}{I_{\text{PAX}}}
+$$
+
+with $I_{\text{PAX}}$ as revenue per seat and flight (see following table).
+
+Segment | $I_{\text{PAX,multi-class}}$ in EUR/SO | $I_{\text{PAX,all-economy}}$ in EUR/SO |
+----------------|:--------------------------------------:|:--------------------------------------:|
+Short haul | 400 | 250 |
+Medium haul | 450 | 300 |
+Long haul | 550 | 400 |
+Ultra long haul | 700 | 550 |
Finally, the SKC correction can be determined as follows:
-$
- SKC_{cor} = SKC \cdot \frac{n_{PAX}}{n_{PAX} + n_{PAX,cargo}}
-$
+$$
+ SKC_{\text{cor}} = SKC \cdot \frac{n_{\text{PAX}}}{n_{\text{PAX}} + n_{\text{PAX,cargo}}}
+$$
-### Ton kilometer costs (calculate_ton_kilometer_costs function)
+### Ton kilometer costs
The ton kilometer costs (TKC) allow the analysis of a change in payload capacity and thus the assessment of the payload kilometer potential. The Ton Kilometers Offered (TKO) are the product of the payload and the range:
-$
- TKO = m_{PL} \cdot R
-$
+$$
+ TKO = m_{\text{PL}} \cdot R
+$$
+
The Ton Kilometer Costs (TKC) are the DOC related to the TKO:
-$
+$$
TKC = \frac{DOC}{TKO}
-$
+$$
-### Revenue seat kilometer costs (calculate_revenue_seat_kilometer_costs)
-Revenue passenger kilometers (RPK) are a measure of how many kilometers the aircraft has carried paying passengers. It is often referred to as "traffic" as it represents the actual demand for air transport. The RPK are determined by multiplying the range by the number of paying passengers. The revenue passenger kilometers are calculated by multiplying the number of revenue passengers with the maximum number of seats and the seat load factor $f_{PL}$:
-$
- RPK = n_{PAX} \cdot f_{SL} \cdot R
-$
+### Revenue seat kilometer costs
+Revenue passenger kilometers (RPK) are a measure of how many kilometers the aircraft has carried paying passengers. It is often referred to as "traffic" as it represents the actual demand for air transport. The RPK are determined by multiplying the range by the number of paying passengers. The revenue passenger kilometers are calculated by multiplying the number of revenue passengers with the maximum number of seats and the seat load factor $f_{\text{SL}}$:
+$$
+ RPK = n_{\text{PAX}} \cdot f_{\text{SL}} \cdot R
+$$
The DOC per revenue passenger kilometer additionally take into account the overall performance of an airline. Note that revenue is strongly dependent on market situation and therefore varying.
-$
+$$
RSKC = \frac{DOC}{RPK}
-$
+$$
## Indirect operating costs (IOC)
tbd. :construction:
diff --git a/docs/documentation/analysis/cost_estimation/run_your_first_cost_estimation.md b/docs/documentation/analysis/cost_estimation/run_your_first_cost_estimation.md
index a043b4a7f388d03eb83461e3370f641e4c3b6dbb..595234bdf9588970ec6ce5e9b88c35f22b3b73ed 100644
--- a/docs/documentation/analysis/cost_estimation/run_your_first_cost_estimation.md
+++ b/docs/documentation/analysis/cost_estimation/run_your_first_cost_estimation.md
@@ -3,13 +3,14 @@ Let's dive into the fun part and crunch some numbers! :moneybag:
## Tool single execution
The tool can be executed from the console directly if all paths are set. The following will happen:
+
- [Console output](#console-output)
- [Generation of reports and plots](#reporting)
- [Writing output to aircraft exchange file](#write-data-to-acxml)
Some of the above mentioned steps did not work? Check out the [troubleshooting](#troubleshooting) section for advices. Also, if you need some additional information on the underlying methodology, check out the page on the [cost estimation method](operating_cost_method.md).
-So, feel free to open the terminal and run `cost_estimation.exe` to see what happens...
+So, feel free to open the terminal and run `python.exe cost_estimation.py` to see what happens...
### Console output {#console-output}
Firstly, you see output in the console window. Let's go through it step by step...
@@ -45,7 +46,7 @@ The tool continues to check if an off-design study exists and tries to calculate
```
2024-12-06 11:37:30,641 - PRINT - Plots are generated and saved...
-2024-12-06 11:37:38,187 - WARNING - Warning: "html_output" switch in module configuration file set to "False". No HTML report generated.
+2024-12-06 11:37:38,187 - PRINT - HTML report is generated and saved...
2024-12-06 11:37:38,188 - PRINT - Method-specific data are written to 'cost_estimation_results.xml'...
2024-12-06 11:37:38,192 - WARNING - Warning: "tex_output" switch in module configuration file set to "False". No TeX report file generated.
2024-12-06 11:37:38,192 - PRINT - Cost estimation finished.
@@ -54,9 +55,10 @@ Finally, you receive information about the reports and plots created (depending
### Reporting {#reporting}
In the following, a short overview is given on the generated reports:
+
- A `cost_estimation.log` file is written within the directory of the executable
- Depending on your settings, the following output is generated and saved in the `reporting` folder, located in the directory of the aircraft exchange file:
- - an HTML report in the `report_html` folder (not implemented yet)
+ - an HTML report in the `report_html` folder
- a TeX report in the `report_tex` folder (not implemented yet)
- an XML file with additional output data in the `report_xml` folder
- plots in the `plots` folder
diff --git a/docs/documentation/analysis/ecological_assessment/basic-concepts.md b/docs/documentation/analysis/ecological_assessment/basic-concepts.md
index cfef150f484dbdb4696fbbfa5348df6b6f39b744..5c787023b5a78a21241b1fd55311ccadb249b21a 100644
--- a/docs/documentation/analysis/ecological_assessment/basic-concepts.md
+++ b/docs/documentation/analysis/ecological_assessment/basic-concepts.md
@@ -574,11 +574,11 @@ with
The forcing factors can be interpolated from given data and average values per mission are determined, considering climb, cruise and approach phase.
-In a next step, the radiative forcing is normalized with the species' efficacy f and $RF_{2xCO_2}$, the RF which would result from a doubling of CO2:
+In a next step, the radiative forcing is normalized with the species' efficacy f and $RF_{2\times CO_2}$, the RF which would result from a doubling of CO2:
-$$
- RF^{*}_{i}(t,h) = f_i \cdot \frac{RF_i(t,h)}{RF_{2xCO_2}}
-$$
+<div class="mathjax-render">
+$ RF^{*}_{i}(t,h) = f_i \cdot \frac{RF_i(t,h)}{RF_{\left(2 \times CO_2\right)}} $
+</div>
With these values, a temperature change can be determined:
@@ -597,12 +597,14 @@ $\Delta T_{weighted}(t) = \Delta T(t) \cdot w(t)$
with the weighting function:
-$w(t) = \begin{cases}
- 1, & t < H \\
- \frac{1}{(1 + r)^{(t - H)}}, & H < t\le t_{max}\\
- 0, & t > t_{max}
- \end{cases}
-$
+<div class="mathjax-render">
+ $ w(t) = \begin{cases}
+ 1, & t < H \\
+ \frac{1}{(1 + r)^{(t - H)}}, & H < t\le t_{max}\\
+ 0, & t > t_{max}
+ \end{cases}
+ $
+</div>
with
@@ -613,7 +615,7 @@ with
In a last step, the average temperature response ATR is calculated:
-$ATR_i = \frac{1}{H} \cdot \int_{0}^{t_{max}} \Delta T_{weighted}(t) $
+$ ATR_i = \frac{1}{H} \cdot \int_{0}^{t_{max}} \Delta T_{weighted}(t) $
The overall ATR is then:
diff --git a/docs/documentation/analysis.md b/docs/documentation/analysis/index.md
similarity index 85%
rename from docs/documentation/analysis.md
rename to docs/documentation/analysis/index.md
index a2581a3bcf8dc3b722b60ea842566216b4f0f83f..2368eb809a9e520755c8524c01819b2eafbf0151 100644
--- a/docs/documentation/analysis.md
+++ b/docs/documentation/analysis/index.md
@@ -12,7 +12,7 @@ glightbox: false
---
## Aerodynamic analysis
-{.overview-img align=left}
+{.overview-img align=left}
The tool `aerodynamic_analysis` calculates, as the tool name suggests, the polars of an aircraft.
It uses the tool Lifting Line from DLR to calculate force, lift and moment coefficients for each lifting surface of the aircraft.
These coefficients are used to calculate induced, viscous and wave drag as well as the moment coefficients for the overall aircraft.
@@ -27,7 +27,7 @@ lift mach numbers.
---
## Mission analysis
-{.overview-img align=left}
+{.overview-img align=left}
The module `mission_analysis` is the key module of the aircraft performance analysis.
Its purpose is to calculate the flight trajectory, based on the inputs of the preliminary aircraft design cycle, by solving the aircraft equations of motion being simplified as a point mass model.
Depending on the method, the fuel consumption is calculated either:
@@ -44,7 +44,7 @@ For the user, possible changes in the module run configuration can be made in th
---
## Weight and balance analysis
-{.overview-img align=left}
+{.overview-img align=left}
The `weight_and_balance_analysis` module calculates sub-masses and total masses of the aircraft including center of gravities.
{.overview-item}
@@ -55,18 +55,18 @@ The `weight_and_balance_analysis` module calculates sub-masses and total masses
---
## Constraint analysis
-{.overview-img align=left}
+{.overview-img align=left}
The `constraint_analysis` module updates the performance criteria wing loading and thrust-to-weight-ratio based on the calculated aircraft data.
{.overview-item}
|Module Version|Language|License|Documentation|
|:---:|:---:|:---:|---|
-|1.0.0|:simple-cplusplus: |GPLv3|-|
+|1.0.0|:simple-cplusplus: |GPLv3|[Link](analysis/constraint_analysis/index.md)|
---
## Ecological assessment
-{.overview-img align=left}
+{.overview-img align=left}
The **ecological_assessment** is an additional module of the UNICADO toolchain.
Its purpose is to calculate the emissions and energy demand within the aircraft's lifecycle and to determine the missions based climate impact.
For the user, possible changes in the module run configuration can be made in the related calculateEmissions_conf.xml file.
@@ -83,7 +83,7 @@ The parameters comprised in this XML file can have different attributes as e.g.
---
## Performance assessment
-{.overview-img align=left}
+{.overview-img align=left}
The module `calculatePerformance` is used to evaluate the mission performance of the design.
{.overview-item}
@@ -94,7 +94,7 @@ The module `calculatePerformance` is used to evaluate the mission performance of
---
## Cost estimation
-{.overview-img align=left}
+{.overview-img align=left}
This modules calculates the direct operating cost (DOC) of an aircraft.
Direct costs include all expenses incurred in operating and financing the aircraft:
diff --git a/docs/documentation/analysis/mission_analysis/figures/acceleration/PP_for_gif.pptx b/docs/documentation/analysis/mission_analysis/figures/acceleration/PP_for_gif.pptx
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diff --git a/docs/documentation/analysis/mission_analysis/figures/acceleration/acceleration.drawio b/docs/documentation/analysis/mission_analysis/figures/acceleration/acceleration.drawio
new file mode 100644
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text-align: left;"><div xmlns="http://www.w3.org/1999/xhtml" style="display: flex; align-items: unsafe center; justify-content: unsafe center; width: 170px; height: 1px; padding-top: 82px; margin-left: 735px;"><div data-drawio-colors="color: rgb(0, 0, 0); " style="box-sizing: border-box; font-size: 0px; text-align: center;"><div style="display: inline-block; font-size: 12px; font-family: Helvetica; color: rgb(0, 0, 0); line-height: 1.2; pointer-events: all; white-space: normal; overflow-wrap: normal;"><font style=""><font style="font-size: 18px; border-color: var(--border-color);">h</font><font style="border-color: var(--border-color);"><font style="border-color: var(--border-color); font-size: 9px;">end, i</font><font style="font-size: 18px; border-color: var(--border-color);"> </font><span style="font-size: 18px;">= h</span><font style="border-color: var(--border-color); font-size: 9px;">end, i-1</font></font><span style="font-size: 18px;">?</span></font></div></div></div></foreignObject><text x="820" y="86" fill="rgb(0, 0, 0)" font-family="Helvetica" font-size="12px" text-anchor="middle">hend, i = hend, i-1?</text></switch></g><rect x="220" y="69" width="160" height="30" rx="4.5" ry="4.5" fill="rgb(255, 255, 255)" stroke="rgb(0, 0, 0)" pointer-events="all"/><g transform="translate(-0.5 -0.5)"><switch><foreignObject pointer-events="none" width="100%" height="100%" requiredFeatures="http://www.w3.org/TR/SVG11/feature#Extensibility" style="overflow: visible; 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+
+# Getting started {#getting_started}
+
+Tickets :ticket: please: We are about to start! In this guide, we will show you how to set up your first mission using our **mission_analysis** tool.
+
+
+## Step-by-step
+
+To be able to execute **mission_analysis**, you have to provide the following data beside your [Aircraft Exchange File](#acxml):
+
+- `mission_data` (e.g. `design_mission.xml`)
+- `aero_data` (polar files)
+- `engine_data` (engine maps)
+
+!!! note
+ Those files are generated by [Create Mission XML](../../sizing/create_mission_xml/index.md), [Aerodynamic Assessment](../../sizing/aerodynamic_analysis/index.md) and [Propulsion Design](../../sizing/propulsion_design/index.md) and shall not be edited manually!
+
+To do so, you can either use:
+
+- a pre-calculated aircraft configuration (e.g. from the `Aircraft References` repository),
+- an aircraft project in which the [Sizing Tools](../../sizing/sizing.md) and the [Aerodynamic Assessment](../../sizing/aerodynamic_analysis/index.md) tool have already been executed at least once.
+
+Once your aircraft is ready, you only need to follow these steps to start your calculation:
+
+1. Head over to `mission_analysis_conf` (more details [here](#config_file)). Assuming this file represents the version of the develop branch, edit the following nodes within `control_settings`:
+ - set `aircraft exchange file_name` and `aircraft exchange file_directory` to your respective settings,
+ - set the `plot_output` to false if you don't have `inkscape` or `gnuplot` installed or define `inkscape_path` and `gnuplot_path` if their directories are not registered in your system environments.
+2. Open your terminal within the `mission_analysis` folder and run the **mission_analysis** executable.
+3. Fasten your seatbelt: We are ready for takeoff! :airplane:
+
+If everything is set up correctly, your first `design_mission` should land a few seconds later :star:
+
+## First iteration results
+
+!!! note
+ If you are using a pre-calculated aircraft, **mission_analysis** will generate its results using parameters from the previous calculations. Therefore, the behavior for an initial execution can not be observed. Continue with [Further Iterations](#further_iterations).
+
+Due to many dependencies between the [sizing tools](../../sizing/sizing.md), performance data and component parameters are quite off within the first iteration. This can lead to an unstable aircraft configuration that will fail the `design_mission` (e.g. wrongly sized engines can't climb to the initial cruise altitude). To avoid this, the [low-fidelity 3D Standard Mission](methods.md/#lowfi) (`design_mission::breguet`) is triggered if no previous mission calculation can be found. Unlike the ordinary mission calculation, this sub-version of the `design_mission` finishes after a rough estimation of the fuel consumption. Once this method is finished, the `masses_cg_inertia/maximum_takeoff_mass/mass_properties/mass` node is updated and this block is written into the [Aircraft Exchange File](#acxml):
+
+```xml
+<mission description="Mission data" tool_level="0">
+ <design_mission description="Data of design mission">
+ <range description="Traveled range from break release to end of taxi at destination">
+ <value>4500000</value>
+ <unit>m</unit>
+ <lower_boundary>0</lower_boundary>
+ <upper_boundary>5000000</upper_boundary>
+ </range>
+ <loaded_mission_energy description="Amount of energy loaded into tanks (including reserves) for the mission">
+ <mission_energy ID="0" description="Amount of energy loaded into tanks (including reserves) for specified energy carrier">
+ <consumed_energy description="Energy amount">
+ <value>7.0e+11</value>
+ <unit>J</unit>
+ <lower_boundary>0</lower_boundary>
+ <upper_boundary>1e+13</upper_boundary>
+ </consumed_energy>
+ <energy_carrier_ID description="See energy carrier specification node">
+ <value>0</value>
+ <unit>1</unit>
+ <lower_boundary>0</lower_boundary>
+ <upper_boundary>5</upper_boundary>
+ </energy_carrier_ID>
+ </mission_energy>
+ </loaded_mission_energy>
+ <in_flight_energy description="Amount of energy needed for in-flight segments (all segments from takeoff to landing)">
+ <trip_energy ID="0" description="Amount of energy needed for trip segments (all segments from takeoff to landing) for specified energy carrier">
+ <consumed_energy description="Energy amount">
+ <value>5.5e+11</value>
+ <unit>J</unit>
+ <lower_boundary>0</lower_boundary>
+ <upper_boundary>1e+13</upper_boundary>
+ </consumed_energy>
+ <energy_carrier_ID description="See energy carrier specification node">
+ <value>0</value>
+ <unit>1</unit>
+ <lower_boundary>0</lower_boundary>
+ <upper_boundary>5</upper_boundary>
+ </energy_carrier_ID>
+ </trip_energy>
+ </in_flight_energy>
+ <taxi_energy description="Amount of energy needed for taxiing specified energy carrier">
+ <taxi_out_energy ID="0" description="Amount of energy needed for taxiing at origin for specified energy carrier">
+ <consumed_energy description="Energy amount">
+ <value>1.0+10</value>
+ <unit>J</unit>
+ <lower_boundary>0</lower_boundary>
+ <upper_boundary>1e+13</upper_boundary>
+ </consumed_energy>
+ <energy_carrier_ID description="See energy carrier specification node">
+ <value>0</value>
+ <unit>1</unit>
+ <lower_boundary>0</lower_boundary>
+ <upper_boundary>5</upper_boundary>
+ </energy_carrier_ID>
+ </taxi_out_energy>
+ <taxi_in_energy ID="0" description="Amount of energy needed for taxiing at destination for specified energy carrier">
+ <consumed_energy description="Energy amount">
+ <value>5.0e9</value>
+ <unit>J</unit>
+ <lower_boundary>0</lower_boundary>
+ <upper_boundary>1e+13</upper_boundary>
+ </consumed_energy>
+ <energy_carrier_ID description="See energy carrier specification node">
+ <value>0</value>
+ <unit>1</unit>
+ <lower_boundary>0</lower_boundary>
+ <upper_boundary>5</upper_boundary>
+ </energy_carrier_ID>
+ </taxi_in_energy>
+ </taxi_energy>
+ </design_mission>
+</mission>
+```
+
+
+## Further iterations {#further_iterations}
+
+After the initial loop, we expect a robuster behavior which we can use to calculate the flight segments with an increased resolution. To achieve this, every segment is split into little time and way increments (only a few seconds/meters per increment) aiming for the trajectory points that were written into the `mission file`. In each increment, all relevant parameters are saved into a `mission profile`. After the calculation is done, said `mission profile` is exported as a [CSV file](#csv_file) into the `mission_data` directory. Within the [Aircraft Exchange File](#acxml) the `masses_cg_inertia/maximum_takeoff_mass/mass_properties/mass` node is updated when calculating a `design_mission`; for a `study_mission` it's the `mission/study_mission/takeoff_mass` node. Having a higher resolution also increases the amount of data in the `mission` block:
+
+
+```xml
+<mission description="Mission data" tool_level="0">
+ <design_mission description="Data of design mission">
+ <range description="Traveled range from break release to end of taxi at destination">
+ <value>4500000</value>
+ <unit>m</unit>
+ <lower_boundary>0</lower_boundary>
+ <upper_boundary>5000000</upper_boundary>
+ </range>
+ <loaded_mission_energy description="Amount of energy loaded into tanks (including reserves) for the mission">
+ <mission_energy ID="0" description="Amount of energy loaded into tanks (including reserves) for specified energy carrier">
+ <consumed_energy description="Energy amount">
+ <value>7.0e+11</value>
+ <unit>J</unit>
+ <lower_boundary>0</lower_boundary>
+ <upper_boundary>1e+13</upper_boundary>
+ </consumed_energy>
+ <energy_carrier_ID description="See energy carrier specification node">
+ <value>0</value>
+ <unit>1</unit>
+ <lower_boundary>0</lower_boundary>
+ <upper_boundary>5</upper_boundary>
+ </energy_carrier_ID>
+ </mission_energy>
+ </loaded_mission_energy>
+ <in_flight_energy description="Amount of energy needed for in-flight segments (all segments from takeoff to landing)">
+ <trip_energy ID="0" description="Amount of energy needed for trip segments (all segments from takeoff to landing) for specified energy carrier">
+ <consumed_energy description="Energy amount">
+ <value>5.5e+11</value>
+ <unit>J</unit>
+ <lower_boundary>0</lower_boundary>
+ <upper_boundary>1e+13</upper_boundary>
+ </consumed_energy>
+ <energy_carrier_ID description="See energy carrier specification node">
+ <value>0</value>
+ <unit>1</unit>
+ <lower_boundary>0</lower_boundary>
+ <upper_boundary>5</upper_boundary>
+ </energy_carrier_ID>
+ </trip_energy>
+ <takeoff_energy ID="0" description="Amount of energy needed for takeoff segment for specified energy carrier">
+ <consumed_energy description="Energy amount">
+ <value>8.9e+9</value>
+ <unit>J</unit>
+ <lower_boundary>0</lower_boundary>
+ <upper_boundary>1e+13</upper_boundary>
+ </consumed_energy>
+ <energy_carrier_ID description="See energy carrier specification node">
+ <value>0</value>
+ <unit>1</unit>
+ <lower_boundary>0</lower_boundary>
+ <upper_boundary>5</upper_boundary>
+ </energy_carrier_ID>
+ </takeoff_energy>
+ <landing_energy ID="0" description="Amount of energy needed for landing segment for specified energy carrier">
+ <consumed_energy description="Energy amount">
+ <value>8.9e+9</value>
+ <unit>J</unit>
+ <lower_boundary>0</lower_boundary>
+ <upper_boundary>1e+13</upper_boundary>
+ </consumed_energy>
+ <energy_carrier_ID description="See energy carrier specification node">
+ <value>0</value>
+ <unit>1</unit>
+ <lower_boundary>0</lower_boundary>
+ <upper_boundary>5</upper_boundary>
+ </energy_carrier_ID>
+ </landing_energy>
+ </in_flight_energy>
+ <taxi_energy description="Amount of energy needed for taxiing specified energy carrier">
+ <taxi_out_energy ID="0" description="Amount of energy needed for taxiing at origin for specified energy carrier">
+ <consumed_energy description="Energy amount">
+ <value>1.0e+10</value>
+ <unit>J</unit>
+ <lower_boundary>0</lower_boundary>
+ <upper_boundary>1e+13</upper_boundary>
+ </consumed_energy>
+ <energy_carrier_ID description="See energy carrier specification node">
+ <value>0</value>
+ <unit>1</unit>
+ <lower_boundary>0</lower_boundary>
+ <upper_boundary>5</upper_boundary>
+ </energy_carrier_ID>
+ </taxi_out_energy>
+ <taxi_in_energy ID="0" description="Amount of energy needed for taxiing at destination for specified energy carrier">
+ <consumed_energy description="Energy amount">
+ <value>5.6e+10</value>
+ <unit>J</unit>
+ <lower_boundary>0</lower_boundary>
+ <upper_boundary>1e+13</upper_boundary>
+ </consumed_energy>
+ <energy_carrier_ID description="See energy carrier specification node">
+ <value>0</value>
+ <unit>1</unit>
+ <lower_boundary>0</lower_boundary>
+ <upper_boundary>5</upper_boundary>
+ </energy_carrier_ID>
+ </taxi_in_energy>
+ </taxi_energy>
+ <block_time description="Block time for the whole mission: Time from break release to end of taxiing after landing">
+ <value>21000.0</value>
+ <unit>s</unit>
+ <lower_boundary>0</lower_boundary>
+ <upper_boundary>45000</upper_boundary>
+ </block_time>
+ <flight_time description="Flight time for the whole mission">
+ <value>20000.0</value>
+ <unit>s</unit>
+ <lower_boundary>0</lower_boundary>
+ <upper_boundary>44500</upper_boundary>
+ </flight_time>
+ <takeoff_engine_derate description="Engine power demand">
+ <value>1</value>
+ <unit>1</unit>
+ <lower_boundary>0</lower_boundary>
+ <upper_boundary>1</upper_boundary>
+ </takeoff_engine_derate>
+ <cruise description="Characteristics of the cruise segment">
+ <average_lift_coefficient description="Lift coefficient CL_average: Arithmetic mean over the entire cruise flight">
+ <value>0.52</value>
+ <unit>1</unit>
+ <lower_boundary>-0.01</lower_boundary>
+ <upper_boundary>1</upper_boundary>
+ </average_lift_coefficient>
+ <minimum_lift_coefficient description="Minimum cruise flight lift coefficient CL_min">
+ <value>0.49</value>
+ <unit>1</unit>
+ <lower_boundary>-0.01</lower_boundary>
+ <upper_boundary>1</upper_boundary>
+ </minimum_lift_coefficient>
+ <maximum_lift_coefficient description="Maximum cruise flight lift coefficient CL_max">
+ <value>0.56</value>
+ <unit>1</unit>
+ <lower_boundary>-0.01</lower_boundary>
+ <upper_boundary>1</upper_boundary>
+ </maximum_lift_coefficient>
+ <top_of_climb_mass description="Total aircraft mass at top of climb (= start of initial cruise altitude (ICA))">
+ <value>77000.0</value>
+ <unit>kg</unit>
+ <lower_boundary>0</lower_boundary>
+ <upper_boundary>150000</upper_boundary>
+ </top_of_climb_mass>
+ <top_of_descend_mass description="Total aircraft mass at top of descend (TOD)">
+ <value>66000.0</value>
+ <unit>kg</unit>
+ <lower_boundary>0</lower_boundary>
+ <upper_boundary>150000</upper_boundary>
+ </top_of_descend_mass>
+ <top_of_climb_range description="Flown range from takeoff to top of climb (= start of initial cruise altitude (ICA))">
+ <value>220000.0</value>
+ <unit>kg</unit>
+ <lower_boundary>0</lower_boundary>
+ <upper_boundary>500000</upper_boundary>
+ </top_of_climb_range>
+ <top_of_descend_range description="Flown range from takeoff to top of descend">
+ <value>4300000.0</value>
+ <unit>kg</unit>
+ <lower_boundary>0</lower_boundary>
+ <upper_boundary>5000000</upper_boundary>
+ </top_of_descend_range>
+ <cruise_steps description="Cruise step information">
+ <cruise_step ID="0" description="Data of a cruise step">
+ <relative_end_of_cruise_step description="End of cruise step relative to total cruise length">
+ <value>0.5</value>
+ <unit>1</unit>
+ <lower_boundary>0</lower_boundary>
+ <upper_boundary>1</upper_boundary>
+ </relative_end_of_cruise_step>
+ <altitude description="Altitude of cruise step">
+ <value>10058.4</value>
+ <unit>m</unit>
+ <lower_boundary>0</lower_boundary>
+ <upper_boundary>15000</upper_boundary>
+ </altitude>
+ </cruise_step>
+ <cruise_step ID="1" description="Data of a cruise step">
+ <relative_end_of_cruise_step description="End of cruise step relative to total cruise length">
+ <value>1</value>
+ <unit>1</unit>
+ <lower_boundary>0</lower_boundary>
+ <upper_boundary>1</upper_boundary>
+ </relative_end_of_cruise_step>
+ <altitude description="Altitude of cruise step">
+ <value>10668.0</value>
+ <unit>m</unit>
+ <lower_boundary>0</lower_boundary>
+ <upper_boundary>15000</upper_boundary>
+ </altitude>
+ </cruise_step>
+ </cruise_steps>
+ </cruise>
+ </design_mission>
+</mission>
+```
+
+## Additional Output
+
+Beside the output written into the [aircraft XML](#acxml), **mission_analysis** generates a few more files you and even other tools can work with
+
+
+### Mission Data CSV {#csv_file}
+
+Remember that nice graph from this tool's [introduction](index.md)? This is a simple visualization of this CSV file we described above. Depending on the amount of engines, used energy carriers and other inputs, the CSV file may differ a bit, but usually you can expect the following parameters there:
+
+- Time [s]
+- Range [m]
+- Altitude [m]
+- FL [100 ft]
+- Mode name [-]
+- Total mass [kg]
+- Energy carrier (ID)
+- Thrust [N]
+- Fuelflow [kg/s]
+- Fuel consumed (kerosene | ID = 0) [kg]
+- Energy consumed (kerosene | ID = 0) [J]
+- Mach [-]
+- CAS [m/s]
+- TAS [m/s]
+- TAS [kts]
+- ROC [fpm]
+- SAR [m/kg]
+- Aero Config [-]
+- C_L [-]
+- L over D [-]
+- Spoiler Factor [-]
+- Reynolds Number [-]
+- Engine Rating [-]
+- Engine N1 (PW1127G-JM | ID = 0) [-]
+- Engine N1 (PW1127G-JM | ID = 1) [-]
+- Shaft power offtake [W]
+- Bleed [kg/s]
+- Angle of attack [deg]
+- Glidepath angle [deg]
+- Incidence angle (stabilizer) [deg]
+
+Beside being a neat dataset to show-off, [Ecological Assessment](../ecological_assessment/index.md) can go through it to calculate the ecological impact of an aircraft flying the displayed mission.
+
+
+### Reporting
+
+If you don't want to edit your data on your own, but need to see some basic characteristics of your mission, you can simply go to the `reporting` directory next to your [Aircraft Exchange File](#acxml). Within `report_html`, we already provide many graphs and useful insights which might come in handy. If something went wrong or you need to know what **mission_analysis** has done in detail, there is also a `.log` file next to your executable in which the shell output is tracked.
+
+
+## Mission Configuration {#configuration}
+
+Now that we have successfully generated our first mission output, let's see how you can tweak our tool a little bit :sunglasses:
+
+
+### Aircraft Exchange File {#acxml}
+
+Within the `requirements_and_specifications` block of the `aircraft_exchange_file`, the following nodes can affect the behavior of **mission_analysis** (descriptions to be found within that file):
+
+```plaintext
+requirements_and_specifications
+└── mission_files
+ ├── design_mission_file
+ ├── study_mission_file
+ ├── requirements_mission_file
+└── design_specification
+ ├── propulsion
+ ├── skinning
+ │ ├── thickness
+ ├── configuration
+ │ ├── tank_definition
+ ├── energy_carriers
+└── requirements
+ ├── top_level_aircraft_requirements
+ │ ├── maximum_structrual_payload_mass
+ │ ├── design_mission
+ │ ├── study_mission
+ │ ├── takeoff_distance
+ │ ├── landing_field_length
+ │ ├── icao_aerodrome_reference_code (once 4D missions are ready)
+ │ ├── flight_envelope
+ │ │ ├── maximum_operating_mach_number
+ │ │ ├── maximum_operating_velocity
+ │ │ ├── maximum_approach_speed
+ │ │ ├── maximum_operating_altitude
+ │ │ ├── maximum_altitude_one_engine_inoperative
+ │ │ ├── climb_or_descend_segment_gradient
+ ├── additional_requirements
+ │ ├── landing_gear
+```
+
+The `mission_files` node simply saves the names of said files. Within `design_specification`, we extract everything from the propulsion system (including tanks) in order to analyze fuel consumption and thrust generation. In the `top_level_aircraft_requirements` node, we can find performance maxima and characteristics for `design_mission` and `study_mission`. The later provide nodes for the mission planning (initial cruise altitude and speed, fuel planning etc.). In `additional_requirements`, the `landing_gear` node tells us with which `friction_coefficient` and `braking_coefficient` our aircraft will be slowed down after touchdown.
+
+
+### Configuration File {#config_file}
+
+The `control_settings` are standardized in UNICADO and will not be described in detail here. The program settings are structured like this (descriptions are in the `mission_analysis_conf.xml`):
+
+```plaintext
+Program Settings
+└── Program Specific
+ ├── Specific Air Range Plot
+ ├── Exit If Fuel Limit Reached
+ │ ├── Enable
+ │ ├── Allowed Relative Overshoot
+ ├── Exit If Maximum Takeoff Mass Limit Reached
+ │ ├── Enable
+ │ ├── Allowed Relative Overshoot
+└── General
+ ├── Fuel Planning
+ │ ├── Fuel Estimation
+ │ │ ├── Fuel Estimation Switch
+ │ │ ├── Joint Aviation Requirements Parameters
+ │ │ │ ├── Contingency Fuel
+ │ │ │ ├── Use Additional Fuel
+ │ │ │ ├── Extra Fuel
+ │ │ ├── Federal Aviation Regulations Parameter
+ │ │ │ ├── Use Additional Fuel
+ │ ├── Fuel Flow Factor Taxiing
+ │ ├── Holding
+ │ │ ├── Holding Mach Number
+ │ │ ├── Holding Altitude
+ │ │ ├── Use Economical Speed
+ ├── Increase Engine Rating During Climb
+ ├── Polar Switch Mission Point
+ │ ├── Polar Switch Selector
+ │ ├── Absolute Range Flown
+ │ ├── Relative Range Flown
+ │ ├── Absolute Time Passed
+ │ ├── Relative Time Passed
+ ├── Glideslope Interception Distance
+ ├── Use Breguet Estimation In Cruise
+ ├── Iterate Top Of descend Mass
+ ├── Landing
+ │ ├── Rotation Time
+ │ ├── Thrust Reverser
+ │ │ ├── Enable
+ │ │ ├── Deactivation Speed
+ │ │ ├── Efficiency
+ │ ├── Runway Exit Speed
+└── Mode
+ ├── Mission Methods
+ │ ├── Fidelity Level
+ │ ├── Mission Type
+ │ ├── Center Of Gravity Method
+ ├── Rate Of Climb Switch
+└── Precision
+ ├── Acceleration Increment
+ ├── Mach Acceleration Increment
+ ├── Altitude Increment
+ ├── Way Increment
+ ├── Specific Air Range Check Increment
+```
+
+
+In the `program_specific` node, you can specify if the specific air range (SAR) is plotted (when plotting is turned on in the `control_Settings`). In addition, you can allow the tool to exceed the maximum takeoff mass and fuel mass during the design loop. This can be useful when operating at extreme conditions where fluctuation above the maxima shall not trigger an exit immediately.
+
+
+In `general` you can decide how the needed fuel is estimated and you can tell **mission_analysis** in which way it shall behave in different flight segments.
+
+
+The `mode` node lets you choose the methods that are applied. Using the keyword `low`/`mid` you will trigger the low-fidelity/mid-fidelity version of the [Standard Mission](methods.md) method. It also has three sub-methods to differentiate between `design_mission`, `study_mission` and `requirements_mission` which can be selected in the `mission_type` node. Please mind that the low-fidelity method only accepts the `design_mission`. The `rate_of_climb_switch` will only affect the [Climb to Ceiling](mission_steps.md/#climb_to_ceiling_subparagraph) step of the `requirements_mission`. With this option, **mission_analysis** calculates the optimum rate of climb towards service ceiling.
+
+
+Finally, in `precision` you can set the parameters which will define the before mentioned increments of your mission profile.
diff --git a/docs/documentation/analysis/mission_analysis/index.md b/docs/documentation/analysis/mission_analysis/index.md
new file mode 100644
index 0000000000000000000000000000000000000000..8fad77ed00dd8ef6bb6fe5fe86f014dd27280a47
--- /dev/null
+++ b/docs/documentation/analysis/mission_analysis/index.md
@@ -0,0 +1,61 @@
+
+# Introduction {#mainpage}
+
+**mission_analysis** is an assessment tool that outputs the flown mission profile, saves characteristic parameters within that profile and checks if performance requirements are met. The following mission types can be analyzed:
+
+- `design_mission`:
+ - Defines the mission for which the aircraft shall be optimized
+ - $MTOM$ is altered during the design process
+ - Exports the `mission profile` as a CSV file
+ - Except $MTOM$, all other results for the [Aircraft Exchange File](getting_started.md/#acxml) are saved in the `design_mission` node
+- `study_mission`:
+ - Calculates off-design missions
+ - Exports a `mission profile` as a CSV file
+ - All results for the [Aircraft Exchange File](getting_started.md/#acxml) are saved in the `study_mission` node
+- `requirements_mission`:
+ - Checks top-level aircraft requirements and possible maxima (like maximum operating altitude)
+ - In the [Aircraft Exchange File](getting_started.md/#acxml) only the `requirement_compliance` block is edited
+
+Mentioned parameters include the energy consumptions which has a high impact on how the aircraft is sized. That's the reason why (unlike many other assessment tools) its `design_mission` calculation takes place within the design loop of our [RCE Workflow](../../../workflow.md).
+
+Once your mission is calculated, you can choose from a wide range of profile data which allows you to further investigate what your aircraft actually does. Here's a little example graph which visualizes the engines' total fuelflow during a `design_mission`:
+
+<p align="center">
+ <img src="figures/mission_profile.png" alt="Mission Profile" width="85%">
+ <br>
+ <em>Visualization of an example mission profile.</em>
+</p>
+
+
+## Quick Overview
+
+| Mission method | mission type | Status |
+|----------------------------------|---------------------------|----------------------------------------|
+| [3D Standard Mission (low-fidelity)](methods.md/#midfi)|`design_mission::breguet`| running :white_check_mark:|
+| [3D Standard Mission (mid-fidelity)](methods.md/#midfi)|`design_mission` | running :white_check_mark:|
+| [3D Standard Mission (mid-fidelity)](methods.md/#midfi)|`study_mission` | running :white_check_mark:|
+| [3D Standard Mission (mid-fidelity)](methods.md/#midfi)|`requirements_mission` | running :white_check_mark:|
+| [4D_trajectory (high-fidelity)](methods.md/#highfi) |None | under development :construction:|
+
+By now, only a [standard (3D) mission method](methods.md/#midfi) is implemented. Its mid-fidelity version can trigger the three missions mentioned above while the low-fidelity sub-version is only used for the `design_mission`. The later is a Breguet-based estimation of the consumed mission fuel and it is triggered automatically if no initial values where given for the `design_mission`. A 4D trajectory mission is also planned, but it is still in the making.
+
+<pre class='mermaid'>
+ graph TD;
+ A[mission_analysis]-->B[design_mission]
+ B-->E[low-fidelity]
+ B-->F[mid-fidelity]
+ B-->G["(high-fidelity)"]
+ A-->C[study_mission]
+ C-->H[mid-fidelity]
+ C-->I["(high-fidelity)"]
+ A-->D[requirements_mission]
+ D-->J[mid-fidelity]
+ D-->K["(high-fidelity)"]
+</pre>
+
+
+## Where to start
+
+If you want a step-by-step guide to start your first calculation, head over to the [Getting Started](getting_started.md) section. We will show you some basic functionalities and how to get your airplane into the air.
+
+Further details about the methods can be found [here](methods.md).
diff --git a/docs/documentation/analysis/mission_analysis/methods.md b/docs/documentation/analysis/mission_analysis/methods.md
new file mode 100644
index 0000000000000000000000000000000000000000..1d1c41bfc861f2bbb8a6cd7e44c98a6aa950c816
--- /dev/null
+++ b/docs/documentation/analysis/mission_analysis/methods.md
@@ -0,0 +1,110 @@
+
+# Mission Methods {#missions}
+
+Depending on computing resources and needed level of detail, we have set up three different approaches to calculate a mission. Okay... it's only two by now, but the third will come for sure! Let's see, what we can find here.
+
+
+## Breguet Estimation (Low Fidelity) {#lowfi}
+
+In this method, the trip fuel mass $ m_{fuel,\,trip} $ (consumed fuel from takeoff until taxi-in) is calculated using the Breguet range equation. To do so, the time needed for climb and cruise are derived from the `mission file`. The approach segment is neglected since its share is rather small and the engines are set to `maximum_continuous` which will overestimate the needed fuel anyway. Up next, it is calculated how much lift $ \overline{L} $, drag $ \overline{D} $, thrust $ \overline{T} $ and fuel massflow $ \overline{\dot{m}}_{fuel} $ are needed on average to reach the top of climb and the end of cruise. After those values are set, the trip fuel mass is computed in the following way:
+
+$ m_{fuel,\,trip} \approx \sum_{i=0}^{n} m_{fuel,\,i} $
+
+$ m_{fuel,\,0} = 0 $
+
+$ m_{fuel,\,i} = (m_{zero\textrm{-}fuel\,mass} + m_{fuel,\,i-1}) \cdot e^{t_{i} \cdot TSFC_i \cdot g \cdot \frac{\overline{D_i}}{\overline{L_i}}} $
+
+$ TSFC_i = \frac{\overline{\dot{m}}_{fuel,\,i}}{\overline{T}_i} $
+
+
+To get the total fuel carried for the mission $ m_{fuel,\,mission} $, taxi-out $ m_{fuel,\,taxi\textrm{-}out} $ and reserve fuel $ m_{fuel,\,reserve} $ are added:
+
+$ m_{fuel,\,mission} = m_{fuel,\,trip} + m_{fuel,\,taxi\textrm{-}out} + m_{fuel,\,reserve} $
+
+Depending on taxiing procedures and reserve fuel methods like JAR or FAR, the fuel quantities may differ (for more details, [click here](#fuel_planning)).
+
+
+
+!!! note
+ The Breguet Estimation is a sub-method of the 3D Standard Mission and therefore uses the same functions for e.g. taxi fuel calculations. When calculating a `design_mission` without having any mission data yet (first loop), the Breguet Estimation is activated automatically to ensure greater robustness. To trigger this method manually, set the [Configuration File's](getting_started.md/#config_file) `fidelity_level` node to `low`.
+
+
+## 3D Standard Mission (Mid Fidelity) {#midfi}
+
+This standard method for the **mission_analysis** tool calculates a `mission profile` that consists of two space dimension plus one time dimension (range, altitude & time). Due to that, it will not handle more complex trajectories like specific flight paths between two airports. To set up its 2D profile, this method derives various target points from [departure, cruise and approach steps](mission_steps.md) stated in the `mission file`. There, every steps' `mode` indicates how _FlightConditions_ and _OperatingConditions_ shall be manipulated to reach those target points. _FlightConditions_ are used to save performance-related values (like true airspeed and current altitude) while _OperatingConditions_ will tell **mission_analysis** in which conditions the aircraft is operated (e.g. high-lift configuration or engine rating).
+
+
+The following `modes` can be found in the steps of the `mission file`:
+
+- `takeoff`
+- `climb`
+- `climb_to_cruise`
+- `climb_to_ceiling`
+- `change_flight_level_constant_ROC`
+- `accelerate`
+- `change_speed`
+- `change_speed_to_CAS`
+- `change_speed_to_Mach`
+- `cruise`
+- `descend_to_approach`
+- `descend`
+- `level_glide_slope_interception`
+- `landing`
+
+
+!!! note
+ Which `mode` is used will be determined by the steps in the `mission file`. If you want to alter them, check out [Create Mission XML](../../sizing/create_mission_xml/index.md).
+
+
+For each step, the start conditions are initialized using the exit data of the previous one (for `TAKEOFF`, an initial step is given manually where most values are set to $0$). Depending on the `mode` different functions are used to change the current conditions of the aircraft iteratively until the required end conditions of the [steps](mission_steps.md) are reached. Since those iterations are split into many small increments, processing the data takes much longer than the [Breguet Estimation](#lowfi). On the upside, this method offers a superior resolution without which a valid analysis of the mission would not be possible. Also, for each increment the relevant flight parameters are saved into a `mission profile` CSV sheet which can be further analyzed.
+
+
+!!! note
+ [The Breguet Estimation](#lowfi) only estimates the aircraft performance using average values for each flight step. Whether the aircraft is able to deliver the needed thrust or lift throughout the whole mission cannot be assured! To get valid mission profiles, always use the [3D Standard Mission](#midfi)! To use this method, set the [Configuration File's](getting_started.md/#config_file) `fidelity_level` node to `mid`.
+
+
+As you might have noticed, we have only discussed the `mission profile` from takeoff until touchdown, but what about taxiing and reserve fuel? Like the others, taxiing steps are described in the [Mission Steps](mission_steps.md/#taxiing) section while fuel planning will be tackled in the following subparagraph.
+
+
+### Fuel Planning Procedures {#fuel_planning}
+
+If you want to set a specific fuel planning procedure, head over to the [Aircraft Exchange File](getting_started.md/#acxml). There, you can select between EASA's fuel planning (_JAR_), FAA's domestic fuel planning (_FAR_DOMESTIC_) and FAA's flag or supplemental fuel planning (_FAR_FLAG_). How the fuel quantities for the different procedures shall be calculated can be changed in the [Configuration File](getting_started.md/#config_file).
+
+_JAR_ consists of:
+
+- Extra fuel:
+ - Fuel mass that shall be carried at the discretion of the captain
+- Alternate Fuel:
+ - Estimated fuel needed to fly the `alternate_distance` (from `mission_file`) on $FL200$
+- Final Reserve Fuel:
+ - $30\,min$ holding at $1500\,ft$ above destination airport at holding speed and ISA-Conditions
+- Additional Fuel:
+ - $15\,min$ holding at $1500\,ft$ above destination airport at holding speed and ISA-Conditions
+- Contingency Fuel using the maximum of the following quantities:
+ - $5\,min$ holding at $1500\,ft$ above destination airport at holding speed and ISA-Conditions
+ - $5\,\%$ of trip-fuel or $3\,\%$ if en-route alternate is available
+
+
+_FAR_DOMESTIC_ consists of:
+
+- Alternate Fuel:
+ - Estimated fuel needed to fly the `alternate_distance` (from `mission_file`) on $FL200$
+- Final Reserve Fuel:
+ - $45\,min$ at mean cruise fuel consumption
+
+
+_FAR_FLAG_ consists of:
+
+- Alternate Fuel:
+ - Estimated fuel needed to fly the `alternate_distance` (from `mission_file`) on $FL200$
+- Final Reserve Fuel:
+ - $30\,min$ holding at $1500\,ft$ above destination airport at holding speed and ISA-Conditions
+- Additional Fuel:
+ - $15\,min$ holding at $1500\,ft$ above destination airport at holding speed and ISA-Conditions
+- Contingency Fuel:
+ - $10\,\%$ of the total required time from brake release (departure airport) to landing (destination airport) at mean cruise fuel consumption
+
+
+## 4D Trajectory (High Fidelity) {#highfi}
+
+Oops, that is not ready yet. An industrious UNICADO coder is probably working on that right now :unicorn:
diff --git a/docs/documentation/analysis/mission_analysis/mission_steps.md b/docs/documentation/analysis/mission_analysis/mission_steps.md
new file mode 100644
index 0000000000000000000000000000000000000000..ee75836a06619e7546c9c86c61a1f2f1dd77fbec
--- /dev/null
+++ b/docs/documentation/analysis/mission_analysis/mission_steps.md
@@ -0,0 +1,232 @@
+# Mission Steps {#mission_steps}
+
+In this section, you will learn how **mission_analysis** interprets the different mission steps from the `mission file`. Beside that, we show you how the taxiing procedures are implemented.
+
+
+## Mission Step Input Parameters
+
+A mission step can consist of the following nodes:
+
+- `configuration`: Aircraft configuration to identify the right polars for aerodynamic calculations (mandatory)
+- `derate`: Thrust derate to (de)throttle the engines during the step (mandatory)
+- `mode`: Defines the mode of the step (mandatory)
+- `rating`: The engine's thrust rating (mandatory)
+- `shaft_power_takeoff_schedule`: Defines the power the engines must provide for the aircraft systems (mandatory)
+- `bleed_air_takeoff` Schedule: Defines bleed air offtakes the engines must provide for the aircraft systems (mandatory)
+- `altitude`: Altitude at the end of this step.
+- `calibrated_airspeed`: Airspeed at the end of this step
+- `mach_number`: Mach number at the end of this step
+- `rate_of_climb_limit`: Maximum rate of climb during this step
+- `flight_management_system`: Indicator if a flight management system is implemented and what its cost index is (`cruise_step` only)
+- `round_to_regular_flight_level`: Rounded flight levels to the multiples of 10 (`cruise_step` only)
+- `auto_select_optimum_flight_level`: Switch to let **mission_analysis** decide what FL is the best (`cruise_step` only)
+- `glide_path`: Angle between glide path and runway (`approach_step` only)
+
+If you need further information about these, please head other to [Create Mission XML](../../sizing/create_mission_xml/index.md).
+
+
+## Step Modes {#step_modes}
+
+In the following paragraphs, we focus on how the steps' `mode` will manipulate the `mission_profile` from start to landing.
+
+
+### Takeoff {#takeoff_subparagraph}
+
+The `takeoff` is composed of ground run (break release until lift-off) and first climb segment to screen height ($35\,ft$). First, the aircraft is accelerated from $ 0\,\frac{m}{s} $ to the lift-off velocity $ v_{LOF} $ utilizing the `acceleration increments` of the [Configuration File](getting_started.md/#config_file). According to [EASA's CS-25 rules](https://www.easa.europa.eu/en/document-library/easy-access-rules/easy-access-rules-large-aeroplanes-cs-25), $ v_{LOF} $ equals $ 110\,\%$ $v_{MU}$ (minimum unstick speed) for aerodynamically limited aircraft and $ 108\,\%$ $v_{MU}$ for geometry limited aircraft. To generalize the $v_{LOF}$ calculation, a more conservative approach has been implemented. Since the minimum safe climb speed at screen height $v_2$ should always be (moderately) greater than the lift-off speed, the following approximation is used (all velocities are calibrated airspeeds):
+
+$$
+v_{LOF} \approx v_2 \geq 1.2 \cdot v_{stall} = 1.2 \cdot 0.94 \cdot v_{stall,\,1g} = 1.128 \cdot v_{stall,\,1g}
+$$
+
+The 1-g stall speed $v_{stall,\,1g}$ is the speed where lift $L$ is equal to the aircraft's weight $ m_{aircraft} \cdot g $ when operating at maximum lift coefficient $C_{L,\,max}$:
+
+
+$$
+L = m_{aircraft} \cdot g = \frac{1}{2} \cdot \rho \cdot v_{stall,\,1g}^2 \cdot C_{L,\,max} \cdot S_{ref}
+\iff v_{stall,\,1g} = \sqrt{\frac{2 \cdot m_{aircraft} \cdot g}{\rho \cdot C_{L,\,max}\cdot S_{ref}}}
+$$
+
+After the aircraft's lift-off, it [climbs with constant speed](#climb_subparagraph) towards screen height to finish this segment.
+
+
+### Accelerate {#accelerate_subparagraph}
+
+`acceleration` segments activate the _change_speed_at_constant_ROC_ function. This mode is usually used for altitudes below $10\,000\,ft$ where the aircraft's speed is increased while retaining a given rate of climb (departure steps). To do so, the speed gap $\Delta v$ between segment start and end is divided into $n$ smaller steps using the [Configuration File's](getting_started.md/#config_file) `acceleration increment`. Then, for $n$ steps the aircraft's velocity is increased using the `acceleration increment`. For each increment, an iterative loop is initiated in which its end altitude is set like this:
+
+
+$$
+h_{end} = h_{start} + \Delta h = h_{start} + \sin(\frac{\gamma}{2 \cdot \overline{a}} \cdot (v_{start}^2 - v_{end}^2))
+$$
+
+before adapting the other _FlightConditions_ using the _set_segment_end_conditions_ function. Once the end altitude within the iteration loop doesn't change anymore, the parameters have converged. Hence, they are saved into the `mission profile` and the next increment will be calculated.
+
+<p align="center">
+ <img src="../figures/acceleration/iteration.gif" alt="Acceleration flow chart" width="95%">
+ <br>
+ <em>Flow chart displaying the iterative pattern to identify the increment's height change.</em>
+</p>
+
+
+### Change Speed {#change_speed_subparagraph}
+
+See [Accelerate](#accelerate_subparagraph). Unlike `accelerate`, `change_speed` uses a (constant) given glide path angle from the `mission file` to derive a rate of climb. Because ATC regulations demand that you can maintain glide path angles between $0°$ and $3°$ at lower altitudes, it is used for deceleration during approach steps below $10\,000\,ft$. .
+
+
+### Change Speed to CAS {#change_speed_to_CAS_subparagraph}
+
+`change_speed_to_CAS` alters the calibrated airspeed while a given rate of climb from the `mission file` is maintained. It's an adaption of [Accelerate](#accelerate_subparagraph) for altitudes between $10\,000\,ft$ and the transition height $h_{transition}$.
+
+
+### Change Speed to Mach {#change_speed_to_Mach_subparagraph}
+
+`change_speed_to_Mach` alters the Mach number while a given rate of climb from the `mission file` is maintained. It's an adaption of [Accelerate](#accelerate_subparagraph) for altitudes above the transition height $h_{transition}$.
+
+
+### Climb {#climb_subparagraph}
+
+The `climb` mode activates the _change_altitude_at_constant_speed_ function.
+To ensure that the aircraft maintains an efficient aerodynamic behavior, the calibrated airspeed is kept constant while the aircraft's altitude is increased/decreased by $\Delta h$. To achieve this, $\Delta h$ is split into $n$ steps by dividing it by the [Configuration File's](getting_started.md/#config_file) `altitude increment`. By default, the minimum rate of climb $ROC$ with which the new altitudes are reached is set to $100\,\frac{ft}{min}$. The actual $ROC$ is calculated using the glide path $\gamma$ while maintaining the total available thrust $T$:
+
+$$
+\gamma = \arcsin \left(\frac{T-D}{g\cdot m_{aircraft}}\right);
+$$
+
+$$
+ROC = \sin(\gamma) \cdot v_{TAS};
+$$
+
+If no maximum $ROC$ is given by the `mission file`, $ROC$ will be taken from the equation above. Else, it is checked if the given $ROC$ limit is exceeded. If this is the case, $ROC$ is set to the maximum while adapting $\gamma$ and $T$ to it. Analogous to [Change Speed](#change_speed_subparagraph), the increment's _FlightConditions_ are looped until $\gamma$ has converged. Afterwards, they are saved into the `mission profile` and the next increment will be calculated.
+
+
+### Climb to Cruise {#climb_to_cruise_subparagraph}
+
+The `climb_to_cruise` mode adapts [Climb](#climb_subparagraph) with the difference that its minimum rate of climb is set to $ 0\,\frac{ft}{min}$. While climbing towards the initial cruise altitude, the air becomes thinner and colder which leads to an increasing Mach number. Once the design cruise Mach number $M_{cruise}$ is exceeded, a constant CAS climb would lead to compressibility effects which could worsen the aircraft's performance. Therefore, the Mach number is kept constant as soon as $M_{cruise}$ is reached. Therefore, $M_{cruise} \approx M_{transition}$.
+
+The altitude at which this occurs is called transition altitude $h_{transition}$ (aka crossover altitude). $h_{transition}$ is defined as the geopotential pressure altitude at which calibrated airspeed and Mach number are representing the same value of true airspeed ($TAS_{Mach} = TAS_{CAS}$). Using the barometric formula, $h_{transition}$ is computed in the following way:
+
+<p align="center">
+ <img src="../figures/transition_altitude.png" alt="Transition Altitude" width="85%">
+ <br>
+ <em>Climb profile at given IAS/MACH Law [1].</em>
+</p>
+
+$$
+h_{transition} = \frac{T_{h=0}}{\frac{\delta T}{\delta h}} \cdot \left(\frac{p_{transition}}{p_{h=0}}\right)^{\frac{R\cdot \frac{\delta T}{\delta h}}{g} - 1}
+$$
+
+$R$ represents the Gas Constant and $g$ the gravitational acceleration. Within the tropopause, the temperature gradient $\frac{\delta T}{\delta h}$ equals $-0.0065\,[K/m]$; above it is defined as $0\,[K/m]$. For $TAS_{Mach}$, you can simply use Mach number $M_{transition}$ and speed of sound $a_{transition}$ which can also be displayed in relation to sea-level conditions:
+
+$$
+TAS_{Mach} = M_{transition}\cdot a_{transition} = M_{transition}\cdot a_{z=0} \cdot\sqrt{\frac{T_{transition}}{T_{z=0}}}
+$$
+
+$TAS_{CAS}$ is computed using isentropic flow equations:
+
+$$
+TAS_{CAS} = a_{h=0} \sqrt{\frac{2}{\kappa - 1} \cdot \frac{\sqrt{T_{transition}}}{T_{h=0}}\cdot\left(\frac{q}{p_{transition}}+1\right)^{\frac{\kappa -1}{\kappa}}-1}
+$$
+
+Where the the stagnation pressure $q$ is derived from the calibrated airspeed:
+
+$$
+CAS = a_{h=0} \sqrt{\frac{2}{\kappa - 1} \cdot \left(\frac{q}{p_{z=0}}+1\right)^{\frac{\kappa -1}{\kappa}}-1}
+$$
+
+
+Finally, the following statement can be derived for the needed pressure ratio characterizing $h_{transition}$:
+
+$$
+\frac{p_{transition}}{p_{h=0}} = \frac{\left(1 + \frac{\kappa-1}{2} \cdot \left(\frac{CAS}{a_{h=0}}\right)^{2} \right)^{\frac{\kappa}{\kappa-1}} - 1}{\left(1 + \frac{\kappa-1}{2} \cdot M_{transition}^{2} \right)^{\frac{\kappa}{\kappa-1}} - 1}
+$$
+
+
+!!! note
+ To determine the cruise range for the [Cruise](#cruise-cruise_subparagraph) segment, the index on the `mission profile` where the aircraft reaches the `initial_cruise_altitude` is saved for later usage.
+
+
+### Climb to Ceiling {#climb_to_ceiling_subparagraph}
+
+This mode should only be used for `requirements missions`! This mode contains four segments:
+
+1. [Climb to Cruise](#climb_to_cruise_subparagraph).
+2. From there, climb to maximum operating altitude with $ROC = 100\,\frac{ft}{min}$ or with a automated maximum rate of climb by turning on the `rate_of_climb_switch`. Either way, the engines are set to `maximum continuous`.
+3. Keep on climbing with $ROC = 100\,\frac{ft}{min}$. Once the engines fail, climb with $ROC = 50\,\frac{ft}{min}$ until they ultimately fail (end altitude = ceiling altitude).
+4. Reset to cruise altitude and repeat step 2 with one engine inoperative.
+
+
+### Change Flight Level {#change_flight_level_constant_ROC_subparagraph}
+
+The `change_flight_level_constant_ROC` segment adapts the [Climb](climb_subparagraph) mode using a minimum rate of climb from the `mission file`. Typically, this option is used in cruise steps to initiate a flight level change. Due to the fact that the cruise altitude usually is way above the transition altitude, the Mach Number is kept constant during this altitude change (see [Climb to Cruise](#climb_to_cruise_subparagraph) for the explanation).
+
+
+### Cruise {#cruise_subparagraph}
+
+In this segment, the aircraft is moved forward with constant speed and $ROC = 0\,\frac{ft}{min}$. How long this `cruise` segment shall last, is determined by the `relative_segment_length` (`mission_file`) which will be applied to the estimated cruise range. To get the latter, the descend range $R_{descend}$ is estimated using the [Breguet method](#lowfi). Then, the afore saved mission segment for reaching `initial_cruise_altitude` (ICA) provides $R_{ICA}$ leading us to the current `cruise` segment's range:
+
+$$
+R_{cruise} = R_{descend} - R_{ICA}
+$$
+
+
+To iterate through this range, it is split into $n$ steps using the [Configuration File's](getting_started.md/#config_file) `way_increment`. Analogous to [Change Speed](#change_speed_subparagraph), the increment's _FlightConditions_ are looped until its consumed fuel mass has converged. Afterwards, they are saved into the `mission profile` and the next increment will be calculated.
+
+
+Even though this mode is not used to climb, the `auto_select_optimum_flight_level` option (see `cruise_steps` in the `mission file`) can be switched on to alter the flight level during `cruise`. If a better specific air range can be obtained on another flight level, **mission_analysis** will test whether the aircraft would consume less fuel there. If this is the case, [Change Flight Level](#change_flight_level_constant_ROC_subparagraph) will take care of the altitude change. Since for regularity reasons discrete flight levels are mandatory, `round_to_regular_flight_level` assures that only permitted altitudes are applied.
+
+!!! note
+ Even if the specific air range of another flight level might be better, a flight level change can cost more fuel than it saves until the end of cruise! Of course, **mission_analysis** is smart enough to take this into account :nerd:
+
+
+### Descend to Approach {#descend_to_approach_subparagraph}
+
+`descend_to_approach` is used to initiate a descend segment from the current cruise altitude towards approach $(10\,000\,ft)$. It uses the same functions the [Climb](#climb_subparagraph) mode does with the difference that its minimum rate of climb is set to $0\,\frac{ft}{min}$. Like in [Climb to Cruise](#climb-climb_subparagraph), the transition altitude $h_{transition}$ will presumably be crossed in this segment. Therefore, the aircraft first descends with a constant Mach number. When $h_{transition}$ is reached, it continues with a constant CAS climb.
+
+!!! note
+ Since the calibrated airspeed won't further decrease below $h_{transition}$, the demanded velocity for the segment's end (segment's `calibrated_airspeed` node in the `mission file`) must be reached before that altitude. If this is not the case, the aircraft will be automatically decelerated by activating a [Change Speed](#change_speed_subparagraph) segment in between.
+
+
+### Descend {#descend_subparagraph}
+
+This mode adapts [Climb](climb_subparagraph) with the difference that its minimum rate of climb is set to $0\,\frac{ft}{min}$ and a glide path angle $\gamma$ is read from the `mission file`. This comes in handy to meet ATC regulations for lower altitudes. Hence, `descend` should be used for approach steps below $10\,000\,ft$.
+
+!!! note
+ After the last descend segment, **mission_analysis** expects the aircraft to be at threshold crossing height ($50\,ft = 15.24\,m$). Otherwise, [Landing](#landing) might cause problems!
+
+
+### Glide Slope Interception {#level_glide_slope_interception_subparagraph}
+
+With `level_glide_slope_interception` the final approach slope is initiated by [cruising](#cruise_subparagraph) at glide slope interception altitude ($3000\,ft$) with constant calibrated airspeed. The distance until the aircraft reaches the interception point $\Delta x$ is derived from the landing glide slope $\gamma$ (usually it's about $3°$), total range $R_{total}$ and the aircraft's current position:
+
+$$
+\Delta x = R_{total} - \frac{h_{current}}{\tan(-\gamma)} - R_{current}
+$$
+
+!!! warning
+ If $\Delta x$ becomes negative, the interception was overflown. This can happen if e.g. the engine produces too much thrust while decelerating or the drag is too low. Either way, **mission_analysis** will try to land the aircraft, but the result may not be ATC conform.
+
+
+### Landing {#landing_subparagraph}
+
+Like [Descend](#descend_subparagraph), the `landing` mode changes the altitude using a constant calibrated airspeed while maintaining a given glide path angle. After touchdown, the aircraft is decelerated to the dedicated taxi speed. Beside the aircraft's brakes, you can also turn on the `thrust_reverser` in the [Configuration File](getting_started.md/#config_file). This may shorten the needed runway length drastically, but you must be sure your engines/aircraft configuration is capable of this.
+
+
+## Taxiing procedures
+
+Unlike the other mission steps, taxi-out and taxi-in are defined in the overall `mission` block within the `mission file`. The taxi fuel consumption for both the origin and destination is determined based on the type of `taxiing_procedure` used. If electric taxiing is used, fuel is only needed for engine warm-up at the origin airport, while no fuel is allocated for taxiing at the destination. The warm-up fuel is calculated using the `engine_warmup_time` $t_{warm\textrm{-}up}$ time and fuelflow rate $\dot{m}_{warm\textrm{-}up}$ which is derived from the [Configuration File's](getting_started.md/#config_file) `fuel_flow_factor_taxiing` which is applied to the engine running in `idle`:
+
+$$
+m_{fuel,\,warm\textrm{-}up} = t_{warm\textrm{-}up} \cdot \dot{m}_{warm\textrm{-}up}
+$$
+
+If electric taxiing is not used, fuel is needed for both origin and destination taxi operations. In this case, the required fuel mass is based on the taxiing time $t_{taxi}$ at each airport (`taxi_time_origin` & `taxi_time_destination`). Analogous to $\dot{m}_{warm\textrm{-}up}$, we get the taxi fuels:
+
+$$
+m_{fuel,\,taxi\textrm{-}out} = t_{taxi\textrm{-}out} \cdot \dot{m}_{taxi\textrm{-}out}
+$$
+
+$$
+m_{fuel,\,taxi\textrm{-}in} = t_{taxi\textrm{-}in} \cdot \dot{m}_{taxi\textrm{-}in}
+$$
+
+!!!node
+ The fuelflow is computed the same way for the three procedures above. Therefore all of these are equal.
diff --git a/docs/documentation/analysis/weight_and_balance_analysis/basic-concepts.md b/docs/documentation/analysis/weight_and_balance_analysis/basic-concepts.md
index 238d4d7263d1b096c287b400e24e1975758688d6..59b5c00034669887ca1f1fd72b01842840d74a37 100644
--- a/docs/documentation/analysis/weight_and_balance_analysis/basic-concepts.md
+++ b/docs/documentation/analysis/weight_and_balance_analysis/basic-concepts.md
@@ -3,8 +3,8 @@ This chapter introduces the definitions and theoretical concepts used in UNICADO
For some calculations there are more available methods. These can be selected by the user in the w&b tool configuration file [_weight\_and\_balance\_analysis\_conf.xml_](usage.md).
-> [!NOTE]
-> In this beta release of UNICADO the w&b analysis module is laid out for the _tube and wing_ configuration of a look-a-like A320. This will be extended in the future to support also a blended wing body configuration.
+!!! note
+ In this beta release of UNICADO the w&b analysis module is laid out for the _tube and wing_ configuration of a look-a-like A320. This will be extended in the future to support also a blended wing body configuration.
## Masses of the Aircraft {#masses}
@@ -13,34 +13,34 @@ Let us start defining the different masses calculated by the tool and how they a
- The **Manufacture Empty Mass (MEM)** is the mass of the aircraft which includes the mass of the main components, i.e. the airframe structure (wing, fuselage, landing gear, empennage, pylons), the propulsion group (nacelles and engines) mass and the fixed equipment mass like the furnishings or the navigation systems.
-> [!NOTE]
-> The tanks don't have an own mass as they are integrated in the main components. Only for the case of additional tanks a mass is added.
+!!! note
+ The tanks don't have an own mass as they are integrated in the main components. Only for the case of additional tanks a mass is added.
- The **Operating Empty Mass (OEM)** represents the mass of the aircraft which includes the crew, all essential operational fluids and all operator-required items and equipment for flight. It coresponds to the MEM with addition of the operator items mass.
- $ OEM = MEM + operator\_items\_mass $
+ $$ OEM = MEM + operator\_items\_mass $$
-> [!NOTE]
-> The operator items are calculated by both the fueselage design and the systems design module.
+!!! note
+ The operator items are calculated by both the fueselage design and the systems design module.
- The **Maximum Zero Fuel Mass (MZFM)** is the total mass of the aircraft without any fuel. It is calculated with
-
- $MZFM = OEM + maximum\_payload\_mass $
+
+ $$ MZFM = OEM + maximum\_payload\_mass $$
- - The ***maximum payload mass*** is refering to the maximum allowed payload which can be taken on board without violation of the structural limits and capacity constraints. This is defined in the TLARs.
+ - The ***maximum payload mass*** is refering to the maximum allowed payload which can be taken on board without violation of the structural limits and capacity constraints. This is defined in the TLARs.
- The **Ferry Range Mass (FRM)** is the mass at which the aircraft can reach the maximum range. For this, no payload is carried and the tanks are filled up with the maximum fuel mass.
- $ FRM = OEM + maximum\_fuel\_mass $
+ $$ FRM = OEM + maximum\_fuel\_mass $$
- - The ***maximum fuel mass*** is the maximum fuel that can be carried and fits in all tanks up to the maximum capacity, i.e all tanks are full. The tank design module outputs the maximum energy per each designed tank. These are transformed here with the corresponding gravimetric density to a maximum fuel mass per tank and then summed up for all tanks.
+ - The ***maximum fuel mass*** is the maximum fuel that can be carried and fits in all tanks up to the maximum capacity, i.e all tanks are full. The tank design module outputs the maximum energy per each designed tank. These are transformed here with the corresponding gravimetric density to a maximum fuel mass per tank and then summed up for all tanks.
- The **Maximum Take-Off Mass (MTOM)** is the mass at which the aircraft takes off. For the design mission this corresponds to the design mass at take-off. Starting with the previously determined OEM, the calculated design fuel at takeoff and the design payload mass are added:
- $ MTOM = OEM + design\_fuel\_mass\_takeoff + design\_payload\_mass $
+ $$ MTOM = OEM + design\_fuel\_mass\_takeoff + design\_payload\_mass $$
-> [!NOTE]
-> The estimated MTOM is an input of the weight and balance analysis tool and is initially written by the _initial\_sizing_ module. Here, it is updated to a mass based on more exact calculation, as the components design and its mass breakdown is now known.
+!!! note
+ The estimated MTOM is an input of the weight and balance analysis tool and is initially written by the _initial\_sizing_ module. Here, it is updated to a mass based on more exact calculation, as the components design and its mass breakdown is now known.
- The ***design payload mass*** consists of the passenger, luggage and additional cargo mass defined by the user in the transport task.
- The ***design fuel mass mission*** is the fuel mass determined from the mission information and is equal to the mission energy (including taxi and reserves) divided by the gravimetric density of the energy provider. If the energy needed to complete the mission is not available or unknown, the design fuel mass is calculated from the difference between the estimated MTOM, OEM and the design payload mass.
@@ -48,21 +48,29 @@ Let us start defining the different masses calculated by the tool and how they a
- The ***design fuel mass midflight*** is calculated by substracting from the design fuel mass at takeoff the fuel consumed during the take-off segment and half of the fuel needed for the cruise segment. These data are provided by the mission module. If not, the design fuel mass midflight is approximated to be half of the design fuel mass at takeoff.
- The ***design fuel mass landing*** corresponds to the remaining fuel in the tanks just after the plane touched down. The minimum fuel mass at landing is determined by substracting from the mission fuel mass the trip fuel mass (containing all flight segments) and the taxi fuel mass before the take-off. If no mission information is available, the minimum design fuel mass at landing is calculated by multiplying the design fuel mass at takeoff with factors for the contingency fuel, alternate fuel and the final fuel reserve.
- With the knowledge about the OEM, the design payload mass and the design fuel masses at different points during flight, the total design masses of the aircraft at specific times can be calculated:
- - ***design mass mission*** (the mass of the aircraft in the parking position before the start) $design\_mass\_mission = OEM + design\_fuel\_mass\_mission + design\_payload\_mass. $
+With the knowledge about the OEM, the design payload mass and the design fuel masses at different points during flight, the total design masses of the aircraft at specific times can be calculated:
+
+ - ***design mass mission*** (the mass of the aircraft in the parking position before the start):
+
+ $$ design\_mass\_mission = OEM + design\_fuel\_mass\_mission + design\_payload\_mass. $$
+
- ***design mass at take-off*** (equal with the MTOM and to the ***design mass*** written in the acxml)
- ***design mass at midflight***
- ***design mass at landing***
-- The **Maximum Landing Mass (MLM)** is the maximum mass at which the pilot of the aircraft is allowed to attempt to land due to structural or other limits.
+The **Maximum Landing Mass (MLM)** is the maximum mass at which the pilot of the aircraft is allowed to attempt to land due to structural or other limits.
Two calculation modes are available:
+
- based on the mission information and the consumed fuel during flight (`default method`):
- $MLM = OEM + design\_fuel\_mass\_landing + design\_payload\_mass $
+
+ $$ MLM = OEM + design\_fuel\_mass\_landing + design\_payload\_mass $$
+
- via the `RWTH regression method`: This calculation uses different formulas depending on whether the maximum takeoff mass exceeds a threshold value of 15,000 kg.
+
1. For Aircraft with *MTOM > 15,000 kg* the following empirical formula is used:
- $MLM = 1.9689 \times MTOM^{0.9248}$
+ $MLM = 1.9689 \times MTOM^{0.9248}$
2. For Aircraft with *MTOM ≤ 15,000 kg* a linear approximation is used:
- $MLM = 0.9009 \times MTOM + 410.85 $
+ $MLM = 0.9009 \times MTOM + 410.85 $
Additionally, two masses are calculated for the case that the aircraft flies either with maximum payload mass or with maximum fuel mass. In both cases the difference up to MTOM is completed with fuel or payload respectively. Based on the loading diagramm, the masses at the most forward and most aft CG positions are also determined.
@@ -71,19 +79,20 @@ Additionally, two masses are calculated for the case that the aircraft flies eit
The knowledge of the center of gravity (CG) position and movement is necessary to ensure the static stability and controllability of the aircraft on the ground and in the air. Based on the results of the detailed mass breakdown of the components with their _mass properties_ information, the total center of gravity of the aircraft can now be determined. The position of the overall CG can generally be determined from the position of the individual centers of gravity w.r.t. a global reference point.
-The calculation involves determining the weighted average of the CG positions for all components. For each axis (_x, y ,z_), the function sums the scaled masses, which are the product of a component’s mass and its CG coordinate for the respective axis. This sum is then divided by the total mass of all components to yield the final CG coordinate for that axis. The global center of gravity ($ \text{CG} $) for a specific axis ($ \text{ax} $) is calculated as:
+The calculation involves determining the weighted average of the CG positions for all components. For each axis (_x, y ,z_), the function sums the scaled masses, which are the product of a component’s mass and its CG coordinate for the respective axis. This sum is then divided by the total mass of all components to yield the final CG coordinate for that axis. The global center of gravity ($CG$) for a specific axis ($ax$) is calculated as:
$
-\text{CG}_{\text{ax}} = \frac{\sum_{i=1}^n (m_i \cdot x_i)}{\sum_{i=1}^n m_i}
+CG_{ax} = \frac{\sum_{i=1}^n (m_i \cdot x_i)}{\sum_{i=1}^n m_i}
$
Where:
+
- $ m_i $ is the mass of the $ i $-th component.
- $ x_i $ is the coordinate of the $ i $-th component along the $ \text{ax} $.
- $ n $ is the total number of components.
-> [!NOTE]
-> It is often common to specify the center of gravity as %MAC.
+!!! note
+ It is often common to specify the center of gravity as %MAC.
### Center of Gravity Shift and the Loading Diagramm
@@ -102,27 +111,31 @@ The loadind diagramm is used to display the permissible range of aircraft mass a
Below is a detailed breakdown of idealized key loading processes and their effects on the CG used to construct the loading diagramm. Given the vast number of possible loading combinations and scenarios, a pre-selection of critical cases—often configuration-dependent— has been made to reduce complexity. The following loading scenarios are considered within UNICADO:
**1. Passenger Boarding**
+
- Critical Scenario: Boarding passengers in a _front-to-rear_ or _rear-to-front_ sequence. These sequences represent extreme cases and can significantly affect the CG position.
- Realistic Scenario: Passengers boarding with free seat selection, typically filling _window seats first, followed by middle and aisle seats_. This simulates common boarding patterns and provides a practical estimation of CG shifts.
**2. Loading of Baggage and Cargo**
+
- For aircraft with similarly sized forward and aft cargo holds, the CG can be deliberately influenced by distributing containers or pallets to achieve a CG favorable for cruise flight. For rear-engine aircraft, the larger cargo hold is typically located forward of the wings. The loading scenario for cargo assumes a symmetric _front-to-rear_ or _rear-to-front_ loading sequence.
**3. Refueling**
+
- Low-/Mid-Wing Aircraft: Fuel is loaded in the following order: inner tank → outer tank → central or fuselage tanks.
- High-Wing Aircraft: Fuel is loaded in reverse: outer tank → inner tank → central or fuselage tanks.
-> [!NOTE]
-> It is assumed that the tanks are filled up symmetrically in the mentioned order up to the maximum capacity of each tank with the fuel mass calculated based on the mission information.
+!!! note
+ It is assumed that the tanks are filled up symmetrically in the mentioned order up to the maximum capacity of each tank with the fuel mass calculated based on the mission information.
**4. Defueling (Fuel Consumption During Flight)**
+
- Low-/Mid-Wing Aircraft:
- Fuel is consumed in the order: central or fuselage tanks → inner tank → outer tank.
- High-Wing Aircraft:
- Fuel is consumed in the reverse order: central or fuselage tanks → inner tank → outer tank.
-> [!NOTE]
-> For the moment only the loading case 3 - 1 - 2 - 4 is implemented. The different selection of the loading scenarios can be made in the _weight\_and\_balance\_analysis\_conf.xml_ file.
+!!! note
+ For the moment only the loading case 3 - 1 - 2 - 4 is implemented. The different selection of the loading scenarios can be made in the _weight\_and\_balance\_analysis\_conf.xml_ file.
Finally, the **most forward and most aft _x_-CG positions** and the corresponding masses are depicted from the resulting diagramm.
@@ -131,12 +144,14 @@ Finally, the **most forward and most aft _x_-CG positions** and the correspondin
Inertia forces arise from the tendency of mass to resist accelerations. For rotational accelerations, these forces are represented by the **mass moment of inertia** terms.These are critical parameters in the analysis and design of aircraft, as they determine the rotational dynamics about the principal axes: roll, pitch, and yaw. These values influence stability, control responsiveness, and handling qualities. The moments of inertia are calculated relative to an axis and depend on the mass distribution of the aircraft. The cross products of inertia (e.g., $ I_{xy} $) arise when the axes are not aligned with the principal axes of the mass distribution.
-In this context the mass moments of inertia about the three principal axes
+In this context the mass moments of inertia about the three principal axes
+
- $ I_{xx} $: About the roll axis
- $ I_{yy} $: About the pitch axis
- $ I_{zz} $: About the yaw axis
-> [!NOTE]
-> The mass moments of inertia are calculated only for the total masses.
+
+!!! note
+ The mass moments of inertia are calculated only for the total masses.
are determined determined by means of the following ***calculation methods:***
@@ -147,7 +162,8 @@ are determined determined by means of the following ***calculation methods:***
- **Pitch**: $I_{yy} = \frac{l^2 M R_y^2}{4} \cdot f_{yy} $
- **Yaw:** $I_{zz} = \frac{\left( \frac{b + l}{2} \right)^2 M R_z^2}{4}$
-Where:
+Where:
+
- $ b $: Wingspan
- $ l $: Fuselage length
- $ M $: Aircraft mass
@@ -194,20 +210,24 @@ Here, $f_{xx}$ and $f_{yy}$ are technology factors set to $0.8$ respectively $0.
#### 3. Using the Component's Inertia
This method involves calculating the total inertia tensor of the aircraft based on its components' individual mass properties. For each inertia component ($I_{xx}$, $I_{xy}$, etc.), the function adds the component's intrinsic inertia and the inertia due to its offset from the reference CG (using the Steiner theorem). The mass moments of inertia are given exemplary
-- around the principal axes ($I_{xx}$, $I_{yy}$,$I_{zz}$):
-$
-I_{xx} = \sum (I_{xx},{\text{component}} + m_{\text{component}} \cdot (p^2 + q^2))
-$
-- around the deviation axes (cross-product terms $I_{xy}$, $I_{xz}$, etc.):
-$
-I_{xy} = \sum (I_{xy},{\text{component}} + m_{\text{component}} \cdot -(p \cdot q))
-$
+- around the principal axes ($I_{xx}$, $I_{yy}$,$I_{zz}$):
+
+ $
+ I_{xx} = \sum (I_{xx,\text{component}} + m_{\text{component}} \cdot (p^2 + q^2))
+ $
+
+- around the deviation axes (cross-product terms $I_{xy}$, $I_{xz}$, etc.):
+
+ $
+ I_{xy} = \sum (I_{xy,\text{component}} + m_{\text{component}} \cdot -(p \cdot q))
+ $
with $p$ and $q$ representing the relative distances between the reference center of gravity (CG) and the current component's CG along the specified axes. Specifically:
+
- $ p $: The distance along the first axis (e.g., x, y, or z).
- $ q $: The distance along the second axis (e.g., x, y, or z).
-> [!NOTE]
-> The component's moments of inertia, if available, are calculated in the component's design modules. Otherwise, these are 0.
+!!! note
+ The component's moments of inertia, if available, are calculated in the component's design modules. Otherwise, these are 0.
diff --git a/docs/documentation/analysis/weight_and_balance_analysis/index.md b/docs/documentation/analysis/weight_and_balance_analysis/index.md
index b251745c8c236e23897499930a66a3a3a07f7e04..b2424fb3aaa266eede919dab75f0dcfbdfe04a5a 100644
--- a/docs/documentation/analysis/weight_and_balance_analysis/index.md
+++ b/docs/documentation/analysis/weight_and_balance_analysis/index.md
@@ -1,5 +1,5 @@
# Introduction {#mainpage}
-The aircraft’s mass plays a crucial role in determining flight performance and evaluating the design, with the ultimate goal being to minimize the operating empty mass. ðŸ‹ï¸â€â™€ï¸ Knowing individual masses is essential for calculating the center of gravity (CG) and determining the placement of critical components like the landing gear and wings. âœˆï¸ The CG significantly affects the aircraft's stability and controllability. An improperly located CG can compromise flight safety, requiring careful planning to ensure it remains within allowable limits throughout the flight, including during fuel consumption and payload variations. This analysis is typically conducted through a weight and balance evaluation using a loading diagram :chart_with_upwards_trend:, which defines the permissible range for combinations of aircraft mass and CG positions. Mass considerations are also fundamental to cost estimation. As an aircraft’s mass increases, it requires more lift, which leads to higher drag, increased thrust demands, elevated fuel consumption, and ultimately greater fuel and operating costs. 💸
+The aircraft’s mass plays a crucial role in determining the flight performance and evaluating the design, with the ultimate goal being to minimize the operating empty mass. ðŸ‹ï¸â€â™€ï¸ Knowing individual masses is essential for calculating the center of gravity (CG) and determining the placement of critical components like the landing gear and wings. âœˆï¸ The CG significantly affects the aircraft's stability and controllability. An improperly located CG can compromise flight safety, requiring careful planning to ensure it remains within allowable limits throughout the flight, including during fuel consumption and payload variations. This analysis is typically conducted through a weight and balance evaluation using a loading diagram :chart_with_upwards_trend:, which defines the permissible range for combinations of aircraft mass and CG positions. Mass considerations are also fundamental to cost estimation. As an aircraft’s mass increases, it requires more lift, which leads to higher drag, increased thrust demands, elevated fuel consumption, and ultimately greater fuel and operating costs. 💸
In UNICADO, the _weight\_and\_balance_analysis_ tool is used to compute the aircraft's masses, determine the CG positions, calculate mass moments of inertia, and generate the loading diagram. The terms "mass" and "weight" are often used interchangeably in aircraft design, though they are scientifically distinct. In this context, both terms are used to refer to the aircraft's mass.
@@ -20,9 +20,9 @@ If you are familiar with these concepts and want to contribute - head over to th
The following pages will help you understand the code structure:
-- [Developer Guide](https://unicado.pages.rwth-aachen.de/unicado.gitlab.io/developer/developer-installation/)
-- [Build Instructions](https://unicado.pages.rwth-aachen.de/unicado.gitlab.io/developer/build/general/)
-- [How to Python in UNICADO](https://unicado.pages.rwth-aachen.de/unicado.gitlab.io/developer/style/python-modularization/)
+- [Developer Guide](../../../get-involved/developer-installation.md)
+- [Build Instructions](../../../get-involved/build/general.md)
+- [How to Python in UNICADO](../../../get-involved/style/python-modularization.md)
- [Weight & Balance Analysis Tool Structure](usage.md)
We appreciate it!
diff --git a/docs/documentation/analysis/weight_and_balance_analysis/usage.md b/docs/documentation/analysis/weight_and_balance_analysis/usage.md
index 117b5484de7edb366213313f994c4222806fea3f..efda403a7fc07ba9357e67df1ba4dc92ddf71b4f 100644
--- a/docs/documentation/analysis/weight_and_balance_analysis/usage.md
+++ b/docs/documentation/analysis/weight_and_balance_analysis/usage.md
@@ -16,15 +16,16 @@ The following requirements are needed for the tool to run:
1. **First**, it is assumed that you have the UNICADO *package* installed including the executables, the database, and the UNICADO *libraries*.
2. As the w&b analysis tool is an analysis tool, the **second requirement** is that the ***sizing modules***, as well that the ***aerodynamic analysis*** and ***mission analysis*** tools were successfully executed beforehand and that the results are written in the Aircraft Exchange File (acXML). The following information must be available (the subcomponents of the required nodes are not listed here):
- - `aircraft_exchange_file/requirements_and_specifications/requirements/top_level_aircraft_requirements`: `maximum_structrual_payload_mass`
- - `aircraft_exchange_file/requirements_and_specifications/design_specification`: `configuration`, `transport_task`, `energy_carriers`
- - `aircraft_exchange_file/component_design` : the `global_reference_point` and the components `wing`, `empennage`, `tank`, `propulsion`, `landing gear`, `systems` each at least with the nodes `position` and `mass_properties`
- - `aircraft_exchange_file/analysis/aerodynamics/reference_values`: `neutral_point`
- - `aircraft_exchange_file/analysis/masses_cg_inertia`: `maximum_takeoff_mass`
- - `aircraft_exchange_file/analysis/mission/design_mission`: `loaded_mission_energy`, `in_flight_energy`, `taxi_energy`
+
+ - `aircraft_exchange_file/requirements_and_specifications/requirements/top_level_aircraft_requirements`: `maximum_structrual_payload_mass`
+ - `aircraft_exchange_file/requirements_and_specifications/design_specification`: `configuration`, `transport_task`, `energy_carriers`
+ - `aircraft_exchange_file/component_design` : the `global_reference_point` and the components `wing`, `empennage`, `tank`, `propulsion`, `landing gear`, `systems` each at least with the nodes `position` and `mass_properties`
+ - `aircraft_exchange_file/analysis/aerodynamics/reference_values`: `neutral_point`
+ - `aircraft_exchange_file/analysis/masses_cg_inertia`: `maximum_takeoff_mass`
+ - `aircraft_exchange_file/analysis/mission/design_mission`: `loaded_mission_energy`, `in_flight_energy`, `taxi_energy`
-> [!NOTE]
-> When the UNICADO workflow is executed the tool is run automatically. In this case, all the required data should be available anyway.
+ !!! note
+ When the UNICADO workflow is executed the tool is run automatically. In this case, all the required data should be available anyway.
3. The `aircraft_exchange_file_name` and `aircraft_exchange_file_directory` are correctly set in the `control settings` part of the _weight\_and\_balance\_analysis\_conf.xml_ file (configXML). The `console_output` should be set at least to `mode_1`.
@@ -45,7 +46,8 @@ ___
A-->I[doc]
</pre>
-@important Since the documentation might be delayed to the development progress - this graph might not have all information yet.
+!!! danger "Important"
+ Since the documentation might be delayed to the development progress - this graph might not have all information yet.
Let's break down the tool structure and see what happens in the most relevant files:
@@ -54,8 +56,10 @@ Let's break down the tool structure and see what happens in the most relevant fi
## Configuration File {#module-configuration-file}
The _weight\_and\_balance\_analysis\_conf.xml_ is structured into two blocks: the control and program settings. The control settings are standardized in UNICADO and will not be described in detail here. But to get started, you have to change at least
+
- the `aircraft_exchange_file_name` and `aircraft_exchange_file_directory` to your respective settings,
- the `console_output` at least to `mode_1`, and
+- the `plot_output` to false (or define `inkscape_path` and `gnuplot_path`).
!!! note
If the tool is executed via the workflow, those settings are set by the workflow settings.
@@ -111,6 +115,7 @@ By changing the program settings im the configXML we can manipulate how the w&b
```
In this part of the configXML we can select the calculation methods and aircraft configuration for the inertia, the maximum landing mass and the modes for the loading scenarios. Each mode has a description and the selection is made by changing the respective `value`. Most of the default modes coming with the package are set to `mode_0`. This means that:
+
- the mass moments of inertia are calculated using the LTH Tables
- the maximum landing mass is calculated based on the mission information and the consumed fuel during flight
- the selected scenario for refueling is to fill up the tanks with the fuel for the design mission
@@ -121,6 +126,7 @@ In this part of the configXML we can select the calculation methods and aircraft
Once the desired methods are selected and the requirements are in place, the tool can run. In order to start the w&b analysis tool, we can execute it directly from the console if all paths are set (see [How to run a tool](howToRunATool.md)) or run the _main.py_ inside the tool folder.
Following will happen:
+
- First, the necessary data and paths are acquired with ***datapreprocessing.py*** and ***usermethoddatapreparation.py***
- Then the ***methodbasic.py*** is executed and you see the output in the console window: The mass properties of the components are first read, then the total masses are calculated, afterwards the cg shift due to refueling, passangers boarding, cargo loading and finally defueling is determined together with the most fwd and aft CG positions
- Next, the calculated data is postprocessed and the outputs are written to the acXML with ***datapostprocessing.py*** and ***usermethoddatapreparation.py***
@@ -191,13 +197,14 @@ The following results are saved in the acXML under `aircraft_exchange_file/analy
</masses_cg_inertia>
```
-> [!TIP]
-> If you are missing some of the terms in here - take a look at [basic concepts](basic-concepts.md).
+!!! tip
+ If you are missing some of the terms in here - take a look at [basic concepts](basic-concepts.md).
---
## Troubleshooting {#trouble}
If the tool does not run properly:
+
- Make sure you have all the paths set up correctly and the specified elements exist
- Go through the log file `weight_and_balance_analysis.txt` and check for warnings and critical messages.
diff --git a/docs/documentation/libraries/engine/index.md b/docs/documentation/libraries/engine/index.md
new file mode 100644
index 0000000000000000000000000000000000000000..dd608dfb201baa055683aa498cbb827ec966bb55
--- /dev/null
+++ b/docs/documentation/libraries/engine/index.md
@@ -0,0 +1,48 @@
+# The `engine` Library in UNICADO
+
+The `_engine_` library serves as the core analysis tool for engine data within UNICADO. It provides access to all possible engine data for every tool in UNICADO. The data can be fixed for an engine or at a given operating point. The data output depends on various factors such as the scale factor and power and bleed offtakes from the engine. The primary objective is to establish a **single source of truth** for engine data retrieval.
+
+## Role in `propulsionDesign`
+Within the `propulsionDesign` module:
+- Engines for the aircraft are selected, and their respective files are copied to the engine directory.
+- The **scale factor** is calculated, determining how the engine's thrust is adjusted to meet aircraft requirements (refer to the `propulsionDesign` documentation).
+
+The `engine` library applies this scale factor, ensuring that aircraft parameters can be accessed without further manual adjustments.
+
+## Engine Data Formats
+The engine data is stored in:
+- `engine.xml` — Contains data **independent** of the operating point.
+- CSV files — Store values **dependent** on:
+
+ - **Mach number**
+ - **Altitude**
+ - **Engine power setting**
+
+> **Note:** The data in these files is **raw and unscaled**. The only modification made in `propulsionDesign` is to the fuel flow CSV file, reflecting user-defined efficiency improvements.
+
+## Functionality of the `engine` Library
+The library is responsible for:
+
+- **Reading engine data**
+- **Applying scaling factors to the data**
+- **Modifying values based on performance-influencing factors like bleed and power offtakes**
+
+### Factors Affecting Engine Performance
+The `engine` library incorporates the following factors, either by default or as optional parameters:
+
+- **Scale factor** from `propulsionDesign`
+- **Temperature variations** (non-ISA standard conditions)
+- **Engine derating**
+- **Bleed air extraction** (for turbofan engines)
+- **Spool shaft offtake** (for turbofan engines)
+
+## How the Library Retrieves Data
+- If data is **not dependent** on the operating point → Uses `engine.xml` in a simple readout.
+- If data is **dependent** on the operating point → Uses CSV files and requires:
+
+ - Mach number
+ - Altitude
+ - Engine power setting (e.g., N1 for turbofan engines)
+
+A **linear interpolation** is performed between existing operating points in the deck when retrieving values from CSV files.
+
diff --git a/docs/documentation/libraries.md b/docs/documentation/libraries/index.md
similarity index 86%
rename from docs/documentation/libraries.md
rename to docs/documentation/libraries/index.md
index b39adfa58ab27e3348919a1f2fc1867b1049e7f6..27130690a4574ea638987213cce2c229020bcc26 100644
--- a/docs/documentation/libraries.md
+++ b/docs/documentation/libraries/index.md
@@ -20,7 +20,7 @@ As mentioned in the [build instructions](../get-involved/build/general.md), we h
Currently, only `aircraftGeometry2` is documented.
## aerodynamics
-{.overview-img align=left}
+{.overview-img align=left}
This library helps with interacting with polar data.
It has helper functions to extract and interpolate data of provided airfoil polars.
{.overview-item}
@@ -32,7 +32,7 @@ It has helper functions to extract and interpolate data of provided airfoil pola
---
## aircraftGeometry2
-{.overview-img align=left}
+{.overview-img align=left}
This library is based on the older [aircraftGeometry](#aircraftgeometry) library and extends it to be more modular.
The modularity and flexibility is achieved by using the high performance [Computational Geometry Algorithms Library](https://www.cgal.org/) also known as **CGAL**.
{.overview-item}
@@ -44,7 +44,7 @@ The modularity and flexibility is achieved by using the high performance [Comput
---
## airfoils
-{.overview-img align=left}
+{.overview-img align=left}
The **airfoils** libary provides a database for different airfoils.
{.overview-item}
@@ -55,7 +55,7 @@ The **airfoils** libary provides a database for different airfoils.
---
## aixml
-{.overview-img align=left}
+{.overview-img align=left}
The **aixml** library is the central library which handles the XML files and data access.
It uses a simple XML library, namely *tinyxml*, to read and parse the XML files.
{.overview-item}
@@ -67,7 +67,7 @@ It uses a simple XML library, namely *tinyxml*, to read and parse the XML files.
---
## atmosphere
-{.overview-img align=left}
+{.overview-img align=left}
The **atmosphere** library provides helper functions to calculate atmospheric properties according to the International Standard Atmosphere (*ISA*).
You can set different atmospheric conditions (e.g. *ISA+25*) and calculate the physical properties of the air at different altitudes.
{.overview-item}
@@ -79,7 +79,7 @@ You can set different atmospheric conditions (e.g. *ISA+25*) and calculate the p
---
## blackboxTest
-{.overview-img align=left}
+{.overview-img align=left}
The **blackboxTest** library provides an interface to run a complete module with different test cases and then checks whether a specific result is calculated or set compared to expected values defined in a `blackBoxTestCases.xml`. The tests are realized with the help of the _googleTest_ framework .
{.overview-item}
@@ -90,7 +90,7 @@ The **blackboxTest** library provides an interface to run a complete module with
---
## engine
-{.overview-img align=left}
+{.overview-img align=left}
This library helps with interacting with engine data.
It has helper functions to extract and interpolate data of provided engine data decks.
The engine decks can originate from different softwaretools as long as they provide the same file format.
@@ -111,7 +111,7 @@ UNICADO currently uses two external libaries as submodules:
---
## liftingLineInterface
-{.overview-img align=left}
+{.overview-img align=left}
This library helps with interacting with results provided by the tools **Lifting Line** from DLR.
It has helper functions to extract and interpolate data of the results from the tool.
{.overview-item}
@@ -123,7 +123,7 @@ It has helper functions to extract and interpolate data of the results from the
---
## moduleBasics
-{.overview-img align=left}
+{.overview-img align=left}
This library provides the basis structure for the modular approach of the **UNICADO** tools.
The tools are intended to follow the *Strategy Design Pattern* to execute at different fidelity levels.
The library gives a template how modules should be structured and gives helpers which can be used to select and implement the different fidelity methods.
@@ -136,7 +136,7 @@ The library gives a template how modules should be structured and gives helpers
---
## pymodulepackage
-{.overview-img align=left}
+{.overview-img align=left}
This library provides standardized UNICADO data preprocessing, run, and postprocessing functions for Python modules.
{.overview-item}
@@ -147,7 +147,7 @@ This library provides standardized UNICADO data preprocessing, run, and postproc
---
## runtimeInfo
-{.overview-img align=left}
+{.overview-img align=left}
This library handles the user interface during the modules execution.
In provides custom output streams, which automatically handle the log files and error outputs according to the configuration files.
{.overview-item}
@@ -159,7 +159,7 @@ In provides custom output streams, which automatically handle the log files and
---
## standardFiles
-{.overview-img align=left}
+{.overview-img align=left}
This library provides file interfaces and interacts with the operating system.
It can handle process execution with a simple interface.
The library can handle *UNIX* and *Windows* systems alike.
@@ -175,7 +175,7 @@ The library can handle *UNIX* and *Windows* systems alike.
---
## svl
-{.overview-img align=left}
+{.overview-img align=left}
The `simple vector library` by Andrew Willmott provides vector and matrix classes.
{.overview-item}
@@ -189,11 +189,11 @@ The `simple vector library` by Andrew Willmott provides vector and matrix classe
---
## unitConversion
-{.overview-img align=left}
+{.overview-img align=left}
The **unitConversion** groups the most commonly used unit in aerospace and let's you convert values from one unit to another.
In addition, it defines some common **constants** which are useful for calculations.
{.overview-item}
|Module Version|Language|License|Documentation|Dependencies|
|:---:|:---:|:---:|---|---|
-|2.1.0|:simple-cplusplus: |GPLv3|-|-|
+|2.1.0|:simple-cplusplus: |GPLv3|-|-|
\ No newline at end of file
diff --git a/docs/documentation/sizing/create_mission_xml/figures/flight_path.png b/docs/documentation/sizing/create_mission_xml/figures/flight_path.png
new file mode 100644
index 0000000000000000000000000000000000000000..1f5c8ff062c20cb839f6a9975adfa73345aa1df5
--- /dev/null
+++ b/docs/documentation/sizing/create_mission_xml/figures/flight_path.png
@@ -0,0 +1,3 @@
+version https://git-lfs.github.com/spec/v1
+oid sha256:2b8a41cfe089727a6116f2252b6755310e498bd380b5ec2b476dec80586998ae
+size 191274
diff --git a/docs/documentation/sizing/create_mission_xml/getting_started.md b/docs/documentation/sizing/create_mission_xml/getting_started.md
new file mode 100644
index 0000000000000000000000000000000000000000..51efce0f3b022a6bb6f7cec0f51c5ac8a5722d48
--- /dev/null
+++ b/docs/documentation/sizing/create_mission_xml/getting_started.md
@@ -0,0 +1,288 @@
+# Getting started
+
+Because **create_mission_xml** only needs the user input from the [aircraft XML's](#acxml) `requirements_and_specifications` block it can operate without other tools being executed first. Only if [Mission Analysis](../../analysis/mission_analysis/index.md) was run in between, **create_mission_xml** would adapt given cruise steps, but it is not needed to generate a functioning `mission file`.
+
+
+## Run Create Mission XML
+
+Sounds easy? It gets better! Since the `mission file` is this tool's sole output no plots or reports must be written and therefore, you can simply ignore those settings within the [Configuration File](#config_file). Once you set the path to your aircraft XML and provide its name, you can simply open a terminal and execute **create_mission_xml**. Et voilà , it's done!
+
+Even though, you could happily head over to the next tool, we should take a look at what happens in detail and how you can manipulate the mission for your needs.
+
+
+## Mission Configuration
+
+Like we mentioned before, **create_mission_xml** only needs input from the [aircraft XML's](#acxml) `requirements_and_specifications` block and it's own [Configuration File](#config_file). So, let's see what we can do here!
+
+
+### Aircraft Exchange File {#acxml}
+
+Since we don't need all information from the `requirements_and_specifications` block, we have filtered it a little bit to only show you the relevant nodes:
+
+```plaintext
+requirements_and_specifications
+└── mission_files
+ ├── design_mission_file
+ ├── study_mission_file
+ ├── requirements_mission_file
+└── design_specification
+ ├── transport_task
+ │ ├── cargo_definition
+ │ │ ├── additional_cargo_mass
+ │ ├── passenger_definition
+ │ │ ├── total_number_passengers
+ │ │ ├── mass_per_passenger
+ │ │ ├── luggage_mass_per_passenger
+└── requirements
+ ├── top_level_aircraft_requirements
+ │ ├── maximum_structrual_payload_mass
+ │ ├── design_mission*
+ │ ├── study_mission*
+ │ ├── flight_envelope
+ │ │ ├── maximum_operating_altitude
+ │ │ ├── maximum_approach_speed
+```
+<em>* including its subnodes.</em>
+
+The `mission_files` node simply saves the names of said files. Within `design_specification`, **create_mission_xml** gets the information about the transport task from which we can derive the needed payload mass. In the `top_level_aircraft_requirements` node, the `maximum_structrual_payload_mass` is checked against the calculated payload. Also, we can find other performance maxima and characteristics (e.g. initial cruise altitude & Mach number) for `design_mission` and `study_mission` there.
+
+
+### Configuration File {#config_file}
+
+Since the control settings are equal for all tool's, we will skip it and focus on the tool-specific `program_settings`:
+
+```plaintext
+program_settings
+├── mission_selector
+├── maximum_operating_mach_number
+│ ├── enable
+│ ├── delta
+├── adapt_climb_speed_schedule
+│ ├── enable
+│ ├── crossover_altitude
+├── climb_thrust_setting
+├── maximum_rate_of_climb
+├── design_mission
+│ ├── output_file_name
+│ ├── terminal_operation_time
+│ ├── takeoff_procedure
+│ ├── approach_procedure
+│ ├── taxi_time_origin
+│ ├── taxi_time_destination
+│ ├── auto_select_initial_cruise_altitude
+│ ├── auto_select_flight_level
+│ ├── round_to_regular_flight_level
+│ ├── auto_climb_altitude_steps
+│ ├── auto_rate_of_climb_steps
+│ ├── alternate_distance
+│ ├── engine_warmup_time
+│ ├── taxiing_procedure
+│ ├── origin_airport
+│ ├── destination_airport
+├── study_mission
+│ ├── copy_mach_number
+│ ├── copy_initial_cruise_altitude
+│ ├── output_file_name
+│ ├── terminal_operation_time
+│ ├── takeoff_procedure
+│ ├── approach_procedure
+│ ├── taxi_time_origin
+│ ├── taxi_time_destination
+│ ├── auto_select_initial_cruise_altitude
+│ ├── auto_select_flight_level
+│ ├── round_to_regular_flight_level
+│ ├── auto_climb_altitude_steps
+│ ├── auto_rate_of_climb_steps
+│ ├── alternate_distance
+│ ├── engine_warmup_time
+│ ├── taxiing_procedure
+│ ├── origin_airport
+│ ├── destination_airport
+```
+
+In this config, you can decide what takeoff and approach procedure you want to use and how the aircraft shall operate at the airport and while cruising. In the `mission_selector`, you can choose if the `mission file` shall be generated for a `design_mission`, `study_mission` or `requirements_mission`. For more details, check the descriptions in `create_mission_xml_conf.xml`.
+
+!!!node
+ `maximum_operating_mach_number` and the nodes starting with `auto` will lead to **mission_analysis** ignoring user input from the aircraft XML. In those cases, the tool will try to find an own optimum.
+
+
+## Output
+
+Like we have already discussed, the output of **create_mission_xml** is the mission_file which generally looks like this:
+
+```xml
+<mission>
+ <range description="Mission range">
+ <value>5000000</value>
+ <unit>m</unit>
+ <lower_boundary>0</lower_boundary>
+ <upper_boundary>100000000</upper_boundary>
+ </range>
+ <payload description="Payload mass">
+ <value>20000</value>
+ <unit>kg</unit>
+ <lower_boundary>0</lower_boundary>
+ <upper_boundary>100000</upper_boundary>
+ </payload>
+ <number_of_pax description="Number of passenger (Mass per PAX = 95 kg)">
+ <value>200</value>
+ <unit>1</unit>
+ <lower_boundary>0</lower_boundary>
+ <upper_boundary>1000</upper_boundary>
+ </number_of_pax>
+ <cargo_mass description="Cargo mass">
+ <value>2000</value>
+ <unit>kg</unit>
+ <lower_boundary>0</lower_boundary>
+ <upper_boundary>100000</upper_boundary>
+ </cargo_mass>
+ <desired_cruise_speed description="Planned cruise Mach number for fuel calculation">
+ <value>0.78</value>
+ <unit>1</unit>
+ <lower_boundary>0</lower_boundary>
+ <upper_boundary>1</upper_boundary>
+ </desired_cruise_speed>
+ <alternate_distance description="Distance from destination to alternate aerodrome">
+ <value>370400.2</value>
+ <unit>m</unit>
+ <lower_boundary>0</lower_boundary>
+ <upper_boundary>10000000</upper_boundary>
+ </alternate_distance>
+ <taxi_time_origin description="Taxi time at departure airport">
+ <value>540</value>
+ <unit>s</unit>
+ <lower_boundary>0</lower_boundary>
+ <upper_boundary>10000</upper_boundary>
+ </taxi_time_origin>
+ <taxi_time_destination description="Taxi time at destination">
+ <value>300</value>
+ <unit>s</unit>
+ <lower_boundary>0</lower_boundary>
+ <upper_boundary>10000</upper_boundary>
+ </taxi_time_destination>
+ <engine_warmup_time description="Running time of the engines before take-off">
+ <value>0</value>
+ <unit>s</unit>
+ <lower_boundary>0</lower_boundary>
+ <upper_boundary>10000</upper_boundary>
+ </engine_warmup_time>
+ <terminal_operation_time description="Time at the terminal for stopovers">
+ <value>1500</value>
+ <unit>s</unit>
+ <lower_boundary>0</lower_boundary>
+ <upper_boundary>10000</upper_boundary>
+ </terminal_operation_time>
+ <taxiing_procedure description="Taxiing procedure for start and landing.">
+ <value>propulsion_taxiing</value>
+ </taxiing_procedure>
+ <departure description="Departure procedure; Additional nodes neded for mode...
+ Takeoff: No additional nodes.
+ climb: End Point Altitude [m] (double).
+ accelerate: Rate of climb [m/s] (double), End point CAS [m/s] (double).">
+ <departure_step ID="0" description="Single departure step">
+ <configuration description="Configuration of the aircraft during this step">
+ <value>e.g. clean</value>
+ </configuration>
+ <derate description="Derate during this step">
+ <value>1</value>
+ <unit>1</unit>
+ <lower_boundary>0</lower_boundary>
+ <upper_boundary>1.5</upper_boundary>
+ </derate>
+ <mode description="Mode during this step">
+ <value>e.g. accelerate</value>
+ </mode>
+ <rating description="Sets thrust rating within climb/acceleration segments to Takeoff, Climb, Maximum continuous, Cruise">
+ <value>e.g. idle</value>
+ </rating>
+ <additional_nodes>...</additional_nodes>
+ </departure_step>
+ </departure>
+ <cruise description="Cruise procedure: Additional nodes needed for mode...
+ change_speed_to_CAS: Rate of climb [m/s] (double), end point CAS [m/s] (double)
+ change_speed_to_Mach: Rate of climb [m/s] (double), Mach [-] (double)
+ climb_to_cruise: End Point Altitude [m] (double), Mach [-] (double)
+ cruise: Range [%] (double, relative distance at the end of the cruise segment without climb and descend)
+ change_flight_level_constant_ROC: Rate of climb [m/s] (double), end Point Altitude [m] (double)
+ descend_to_approach: End Point Altitude [m] (double), end point CAS [m/s] (double).">
+ <cruise_step ID="0" description="Single cruise step">
+ <configuration description="Configuration of the aircraft during this step">
+ <value>e.g. clean</value>
+ </configuration>
+ <derate description="Derate during this step">
+ <value>1</value>
+ <unit>1</unit>
+ <lower_boundary>0</lower_boundary>
+ <upper_boundary>1.5</upper_boundary>
+ </derate>
+ <mode description="Mode during this step">
+ <value>e.g. change_speed_to_CAS</value>
+ </mode>
+ <rating description="Sets thrust rating within climb/acceleration segments to Takeoff, Climb, Maximum continuous, Cruise">
+ <value>e.g. idle</value>
+ </rating>
+ <auto_select_optimum_flight_level description="Parameters to handle automatized flight_level changes">
+ <enabled description="Switch for automatic selection of the optimum flight level of the cruise step">
+ <value>false</value>
+ </enabled>
+ <auto_climb_step_height description="Height difference for an automatic altitude change step.">
+ <value>0</value>
+ <unit>m</unit>
+ <lower_boundary>0</lower_boundary>
+ <upper_boundary>5000</upper_boundary>
+ </auto_climb_step_height>
+ </auto_select_optimum_flight_level>
+ <flight_management_system description="Flight management system settings">
+ <enabled description="Switch to indicate if a flight management system is equipped">
+ <value>false</value>
+ </enabled>
+ <cost_index description="Cost index [kg/min], scaled 0 to 999 according to Sperry/Honeywell">
+ <value>0</value>
+ <unit>1</unit>
+ <lower_boundary>0</lower_boundary>
+ <upper_boundary>999</upper_boundary>
+ </cost_index>
+ </flight_management_system>
+ <additional_nodes>...</additional_nodes>
+ </cruise_step>
+ </cruise>
+ <approach description="Approach procedure: Additional nodes needed for mode... :
+ descend: End Point Altitude [m] (double), glide_path [deg] (double).
+ change_speed: End point CAS [m/s] (double), glide_path [deg] (double).
+ level_glide_slope_interception: No additional nodes.
+ landing: End Point Altitude [m] (double), glide_path [deg] (double).">
+ <approach_step ID="0" description="Single approach step">
+ <configuration description="Configuration of the aircraft during this step">
+ <value>e.g. clean</value>
+ </configuration>
+ <derate description="Derate during this step">
+ <value>1</value>
+ <unit>1</unit>
+ <lower_boundary>0</lower_boundary>
+ <upper_boundary>1.5</upper_boundary>
+ </derate>
+ <mode description="Mode during this step">
+ <value>e.g. change_speed</value>
+ </mode>
+ <rating description="Sets thrust rating within climb/acceleration segments to Takeoff, Climb, Maximum continuous, Cruise">
+ <value>e.g. idle</value>
+ </rating>
+ <additional_nodes>...</additional_nodes>
+ </approach_step>
+ </approach>
+</mission>
+```
+
+!!!node
+ Bleed air and power offtakes are not displayed here, but every step will include these, too. Offtakes are written and explained by [Systems Design](../systems_design/index.md).
+
+
+While the most parameters like `range` and `alternate_distance` are copied directly from [Aircraft Exchange File](#acxml) and [Configuration File](#config_file), the `payload` is derived from the given number of passengers, their luggage and the mass per passenger. Each step (`departure_step`, `cruise_step` or `approach_step`) contains the nodes `configuration`, `mode`, `derate` and `rating`. The `configuration` node will tell [Mission Analysis](../../analysis/mission_analysis/index.md) which polar (generated by [Aerodynamic Assessment](../../sizing/aerodynamic_analysis/index.md)) shall be used. `derate` and `rating` characterize the engine operations and `mode` specifies what shall happen during the segment between two steps (more infos about `modes`, [click here](../../analysis/mission_analysis/mission_steps.md/#step_modes)). Furthermore, `cruise_steps` always include `flight_management_system` and `auto_select_optimum_flight_level` nodes.
+
+Other entries within these steps can differ depending on which `mode` is used. What input nodes are needed can be found in the descriptions of `departure`, `cruise` and `approach`. As a rule of thumb, the following input nodes can usually be expected:
+
+- `mode` that changes speed: Target speed (Mach or CAS), rate of climb or target speed
+- `mode` that changes altitude: Target altitude, rate of climb or target speed
+
+!!!node
+ For an `approach_step`, the rate of climb cannot be determined up-front, because the glide path angle must be kept constant at $3°$ due to regulatory requirements. Therefore, the rate of climb will be derived from the `glide_path` node by [Mission Analysis](../../analysis/mission_analysis/index.md).
diff --git a/docs/documentation/sizing/create_mission_xml/index.md b/docs/documentation/sizing/create_mission_xml/index.md
new file mode 100644
index 0000000000000000000000000000000000000000..33892231d6d72779104ea28d16b2123c3deb5c08
--- /dev/null
+++ b/docs/documentation/sizing/create_mission_xml/index.md
@@ -0,0 +1,20 @@
+# Introduction
+
+Good news first: **create_mission_xml** is quite slim... or perhaps the slimmest tool of the whole UNICADO chain. It's sole purpose is to define some basic parameters and target points on the mission's trajectory. Nonetheless, this is critical to the whole operation, because we all know *if you fail planning, you'll be planning your failure!* But no worries, we'll help you out :wink:
+
+
+## What a Mission Looks Like {#typical_mission}
+
+In short, a mission contains a handful of so-called segments with which you can define a basic mission profile. Depending on the aircraft size, regulation and flight path planning philosophy, some details may differ, but in general it should look something like this:
+
+<p align="center">
+ <img src="figures/flight_path.png" alt="Flight segments" width="97.5%">
+ <br>
+ <em>Flight segments with typical speeds: IAS (blue), Mach (green), and TAS (violet) [1].</em>
+</p>
+
+
+**create_mission_xml** sets the target/end points of these flight segments which will later be connected by [Mission Analysis](../../analysis/mission_analysis/index.md). Those target points are saved into the `mission file` in which they are categorized as `departure_steps`, `cruise_steps` and `approach_steps`. What this `mission_file` contains in detail can be found in the [Getting Started](getting_started.md).
+
+
+To fill and order `departure_steps` and `approach_steps`, departure and approach procedures based on regulatory requirements were implemented. Since cruise segments are pretty much straight forward and automatic/optimized flight level changes will be handled by [Mission Analysis](../../analysis/mission_analysis/index.md), there are no pre-defined procedures for cruise. Still, there are a few thing to take into account which we will describe in the [Mission Steps](mission_steps.md) section.
diff --git a/docs/documentation/sizing/create_mission_xml/mission_steps.md b/docs/documentation/sizing/create_mission_xml/mission_steps.md
new file mode 100644
index 0000000000000000000000000000000000000000..27700c5098f298b94cf5ff4cbb346d9551544e6d
--- /dev/null
+++ b/docs/documentation/sizing/create_mission_xml/mission_steps.md
@@ -0,0 +1,126 @@
+# Mission Steps
+
+The steps in the `mission_file` can be filled and arranged in different ways, depending on what departure and approach procedures you want to implement and how you want the aircraft to behave during cruise. Let's see what **create_mission_xml** can do with the `departure`, `cruise` and `approach` nodes.
+
+
+## Departure
+
+| Procedure | Description | Status |
+|---------------------------|----------------------------------------------|---------------------------------|
+| Standard | FAA/ICAO compliant standard | running :white_check_mark: |
+| Standard 19 seat commuter | FAA/ICAO compliant standard 19 seat commuter | running :white_check_mark: |
+| ICAO-A | Noise reduced takeoff according to ICAO | under development :construction:|
+| ICAO-B | Noise reduced takeoff according to ICAO | under development :construction:|
+
+
+### Standard
+
+| Mode | Thrust Rating | Config | End Altitude [m] | Rate of Climb [m/s] | End CAS [m/s] |
+|------------|----------------------|--------------------------------|------------------|-------------------------|-----------------------|
+| takeoff | takeoff | takeoff | N/A | N/A | N/A |
+| climb | climb thrust setting | takeoff landing gear retracted | 457.2 (1500 ft) | maximum rate of climb | N/A |
+| climb | climb thrust setting | takeoff landing gear retracted | 914.4 (3000 ft) | maximum rate of climb | N/A |
+| accelerate | climb thrust setting | climb | N/A | 5.08 (1000 fpm) | 108.0 (210 kt) |
+| accelerate | climb thrust setting | clean | N/A | 6.10 (1200 fpm) | CAS ATC limit climb |
+| climb | climb thrust setting | clean | 3048 (10,000 ft) | maximum rate of climb | N/A |
+
+In the standard procedure, we assume that the thrust-to-weight ratio is high enough to maintain minimum safe climb speed $v_2$ (see [What a Mission Looks Like](index.md/#typical_mission)) from takeoff until en-route transition (`climb` configuration) at $3\,000\,ft$. Please mind, that EASA's CS-25 only allows extrapolation of the propulsion system's takeoff performance data up to that altitude. To do so, the aircraft shall climb with the given `maximum_rate_of_climb` and `climb_thrust_setting` from the [Configuration File](getting_started.md/#config_file) without an acceleration in between. Since the landing gear gets retracted between screen height ($35\,ft$) and $1\,500\,ft$, climbing up to $3\,000\,ft$ is divided into two segments. Like this, it's easier for [Systems Design](../systems_design/index.md) to simulate the retraction and to put the power/bleed air demand into the `mission file`. Once en-route transition is reached, flaps are set to `climb` while accelerating to $210\,kt$ calibrated airspeed. Just after that, the aircraft accelerates further in `clean` configuration (least drag) until the _CAS_ATC_limit_climb_ is obtained. Since the air space below $10,000\,ft$ is more crowded, institutions like FAA and ICAO limit the speed to $250 kt$ calibrated airspeed, but you can change that in the `climb_speed_below_FL100` node of our [Aircraft Exchange File](getting_started.md/#acxml). Then, the aircraft finishes the departure procedure by climbing up to $10,000\,ft$ using the `maximum_rate_of_climb`.
+
+!!!node
+ Although `maximum_rate_of_climb` can be set as a constant value, we usually set it to $-1$ to indicate that the aircraft shall use all possible thrust of its current engine settings to achieve altitude gains. Therefore, rate of climb varies within these climb segments. Since acceleration is most effective and saver when keeping a constant rate of climb, it is manually set to $1\,000\,\frac{ft}{min}$/$1\,200\,\frac{ft}{min}$ which follows the ICAO's recommendations.
+
+
+### Standard 19 seat commuter
+
+| Mode | Thrust Rating | Config | End Altitude [m] | Rate of Climb [m/s] | End CAS [m/s] |
+|--------------|----------------------|--------------------------------|------------------|-------------------------|---------------------|
+| takeoff | takeoff | takeoff | N/A | N/A | N/A |
+| climb | climb thrust setting | takeoff landing gear retracted | 60.96 (200 ft) | maximum rate of climb | N/A |
+| accelerate | climb thrust setting | clean | N/A | 6.10 (1200 fpm) | CAS ATC limit climb |
+| climb | climb thrust setting | clean | 3048 (10000 ft) | maximum rate of climb | N/A |
+
+Using a smaller aircraft, an acceleration segment at lower altitudes is needed. Analogous to the procedure above, it accelerates to the maximum allowed speed (normally $250\,kts$ calibrated airspeed) before climbing with maximum rate of climb towards $10\,000\,ft$.
+
+
+### Minimal Noise Takeoff
+
+ICAO-A and ICAO-B should tackle this, but it is not ready yet :construction:
+
+
+## Cruise
+
+| Mode | Thrust Rating | Config | End Altitude [m] | Rate of Climb [m/s] | End CAS [m/s] / Mach [-] |
+|----------------------|----------------------|---------|------------------|---------------------------|-----------------------------------|
+| change speed to CAS | climb thrust setting | clean | N/A | 1.524 (300 ft/min) | CAS over flight level 100 climb |
+| climb to cruise | climb thrust setting | clean | initial cruise altitude | maximum rate of climb | N/A |
+| change speed to Mach | climb thrust setting | clean | N/A | 0 | initial_cruise_mach_number |
+| cruise | cruise | clean | N/A | N/A | N/A |
+| change flight level constant_ROC / change flight level | cruise | clean | auto / cruise FL + 20 | N/A / auto | N/A |
+| cruise | cruise | clean | N/A | N/A | N/A |
+| descend to approach | idle | clean | 10000 ft | N/A | CAS over flight level 100 descend |
+
+After reaching $10\,000\,ft$ the aircraft accelerates to the next higher speed limit `CAS_over_flight_level_100_climb` which is usually $300\,kts$ calibrated airspeed. Again, you can change this in the aircraft XML, but when you want to stick to current regulations, you should keep $300\,kts (= 154.3334 m/s)$. Then the aircraft keeps on climbing until the `initial_cruise_altitude` from where it accelerates to the `initial_cruise_mach_number` without climbing any further. In the table above, only one flight level change is displayed. How many of them will be initiated can be determined in the following way:
+
+- Short Range ($\leq 1\,000\,NM$):
+ - 1 cruise climb step
+- Medium Range ($1\,000 - 5\,000\,NM$):
+ - 2 cruise climb steps
+- Long Range ($\ge 5\,000\,NM$):
+ - 3 cruise climb steps
+
+!!! node
+ If climbs during cruise are disabled (`no_steps` node in the [Configuration File](getting_started.md/#config_file)), then only 1 climb step is generated. Also when automatic flight level changes are activated, [Mission Analysis](../../analysis/mission_analysis/index.md) will try to find an optimum by itself.
+
+Once the end of cruise is reached, the aircraft shall descend to approach ($10\,000\,ft$) using the maximum descend speed.
+
+!!!node
+ For the `requirements_mission`, `climb_to_cruise` gets replaced by `climb_to_ceiling` where [Missionis Analysis](../../analysis/mission_analysis/index.md) searches for the maximum altitude. After this segment, the mission ends. Thus, the `mission_file` will not have any more entries after `climb_to_ceiling`.
+
+
+## Approach
+
+| Procedure | Description | Status |
+|---------------------------|----------------------------------------------|---------------------------------|
+| Standard | FAA/ICAO compliant standard | running :white_check_mark: |
+| Standard 19 seat commuter | FAA/ICAO compliant standard 19 seat commuter | running :white_check_mark: |
+| Continuous ICAO | Continuous descent approach | under development :construction:|
+| Steep Continuous ICAO | Steep Continuous descent approach | under development :construction:|
+
+
+### Standard
+
+| Mode | Thrust rating | Config | End Altitude [m] | Glide Path Angle [°] | End CAS [m/s] |
+|--------------------------------|----------------|-------------------------------|------------------|----------------------|-------------------------------|
+| change speed | idle | clean | N/A | 0 | CAS ATC limit descend |
+| descend | cruise | clean | 914.4 (3000 ft) | -3 | N/A |
+| change speed | idle | approach | N/A | 0 | $v_{approach}$ |
+| level glide slope interception | cruise | approach landing gear out | N/A | -3 | N/A |
+| change speed | idle | approach landing gear out | N/A | -3 | $v_{max, approach + 5\,kt}$ |
+| descend | cruise | approach landing gear out | 457.2 (1500 ft) | -3 | N/A |
+| change speed | idle | landing | N/A | -3 | $v_{max, approach}$ |
+| descend | cruise | landing | 304.8 (1000 ft) | -3 | N/A |
+| descend | cruise | landing | 15.24 (50 ft) | -3 | N/A |
+| landing | takeoff | landing | 0 | -3 | N/A |
+
+The first approach segment starts at $10\,000\,ft$ where the descend speed limit from the [Aircraft Exchange File](getting_started.md/#acxml) has to be followed. Like in departure, ICAO and FAA dictate $250 kt$ calibrated airspeed. Once this speed limit is met, the aircraft descends to $3,000\,ft$ maintaining a glide path angle of $-3°$. There, high-lift systems are activated (`approach` config) while decelerating to $v_{approach}$. With $v_{approach}$ the aircraft extends its landing gear and cruises to glide slope interception where instrument landing systems start operating. After decelerating to $v_{max, approach + 5\,kt}$, the aircraft descends to visual approach at $1\,500\,ft$. Lastly, the aircraft changes its speed to $v_{max, approach}$ being in `landing` configuration. Tensions rise, while we descend lower and lower until we finally touch the ground. Congratulations, we have landed! You need more braking power? We set the engines' rating to `takeoff` so you can use them as thrust reversers.
+
+!!!node
+ The `maximum_approach_speed` $v_{max, approach}$ (to be found in the [Aircraft Exchange File](getting_started.md/#acxml)) limits the calibrated airspeed below $1,000\,ft$. Above that, $v_{max, approach + 5\,kts} = v_{approach} + 5\,kt$ and $v_{approach} = max\left(v_{max,\,approach + 5\,kts}, 170\,kts\right)$.
+
+
+### Standard 19 seat commuter
+
+| mode | rating | config | End Altitude [m] | Glide Path Angle [°] | End CAS [m/s] |
+|--------------|---------|----------|------------------|----------------------|-----------------------|
+| change speed | idle | clean | N/A | 0 | CAS ATC limit descend |
+| descend | cruise | clean | 609.6 (2000 ft) | -3 | N/A |
+| change speed | idle | approach | N/A | -3 | $v_{max, approach}$ |
+| descend | cruise | landing | 15.24 (50 ft) | -3 | N/A |
+| landing | takeoff | landing | 0 | -3 | N/A |
+
+For smaller aircraft, the approach procedure becomes less complicated. You can simply decelerate to the before mentioned CAS limit of $250\,kt$ before descending towards initial approach fix at $2\,000\,ft$. Next, the aircraft's configuration is set to `approach` while decelerating to $v_{max, approach}$ with which we bring it to the ground using its `landing` configuration. Easy peasy lemon squeezy! :lemon:
+
+
+### (Steep) Continuous Descent Approach
+
+Continuous descent has not been implemented yet, but that's just a matter of time :clock:
diff --git a/docs/documentation/sizing/fuselage_design/design_method.md b/docs/documentation/sizing/fuselage_design/design_method.md
index fdd1ec8dbbb269d8adb955d7a3cc3eece2732f9f..045b78e0b758df4264f85faf78beaced90add548 100644
--- a/docs/documentation/sizing/fuselage_design/design_method.md
+++ b/docs/documentation/sizing/fuselage_design/design_method.md
@@ -13,75 +13,78 @@
The cabin width is estimated using the given class definition.
#### Determine width of seat row per aircraft side
-The width of one seat row/bench $w_{seat\_bench}$ (in inch) can be determined for the left and right side of the aircraft using the following equation:
-$
- w_{seat\_bench} = n_{seats} \cdot w_{seat} + 2 \cdot w_{armrest}
-$
+The width of one seat row/bench $w_{\text{bench}}$ (in inch) can be determined for the left and right side of the aircraft using the following equation:
+$$
+ w_{\text{bench}} = n_{\text{seats}} \cdot w_{\text{seat}} + 2 \cdot w_{\text{armrest}}
+$$
In which
-- $n_{seats}$ - number of seats per seat bench
-- $w_{seat}$ - seat width (taken from lowest class seat)
-- $w_{armrest}$ - armrest width (taken from lowest class seat)
+
+- $n_{\text{seats}}$ - number of seats per seat bench
+- $w_{\text{seat}}$ - seat width (taken from lowest class seat)
+- $w_{\text{armrest}}$ - armrest width (taken from lowest class seat)
#### Calculate cabin width
-The cabin width $w_{cabin}$ (in inch) can then be calculated:
-$
- w_{cabin} = w_{aisle} + w_{seat\_bench\_left} + w_{seat\_bench\_right} + 2 \cdot w_{seat\_space}
-$
+The cabin width $w_{\text{cabin}}$ (in inch) can then be calculated:
+$$
+ w_{\text{cabin}} = w_{\text{aisle}} + w_{\text{bench,left}} + w_{\text{bench,right}} + 2 \cdot w_{\text{seat space}}
+$$
In which
-- $w_{aisle}$ - passenger aisle width
-- $w_{seat\_space}$ - lowest class seat space
-In case of a **wide-body aircraft configuration** there is an additional row in the middle of the aircraft as well as an additional passenger aisle. The width of the seat bench $w_{seat\_bench\_center}$ can be calculated using an equation similar to that in the previous section.
-$
- w_{seat\_bench\_center} = n_{seats} \cdot w_{seat} + 2 \cdot w_{armrest\_outer} + (n_{seats} - 1) \cdot w_{armrest\_inner}
-$
+- $w_{\text{aisle}}$ - passenger aisle width
+- $w_{\text{seat space}}$ - lowest class seat space
+
+In case of a **wide-body aircraft configuration** there is an additional row in the middle of the aircraft as well as an additional passenger aisle. The width of the seat bench $w_{\text{bench,center}}$ can be calculated using an equation similar to that in the previous section.
+$$
+ w_{\text{bench,center}} = n_{\text{seats}} \cdot w_{\text{seat}} + 2 \cdot w_{\text{armrest,outer}} + (n_{\text{seats}} - 1) \cdot w_{\text{armrest,inner}}
+$$
In which
-- $w_{seat}$ - seat width (from lowest class seat parameters of right side)
-- $w_{armrest\_outer}$ - width of outer armrest (from lowest class seat parameters of right side)
-- $w_{armrest\_inner}$ - width of inner armrest (from lowest class seat parameters of right side)
+
+- $w_{\text{seat}}$ - seat width (from lowest class seat parameters of right side)
+- $w_{\text{armrest,outer}}$ - width of outer armrest (from lowest class seat parameters of right side)
+- $w_{\text{armrest,inner}}$ - width of inner armrest (from lowest class seat parameters of right side)
The equation for the cabin width estimation must be adapted accordingly:
-$
- w_{cabin} = w_{aisle} + w_{seat\_bench\_left} + w_{seat\_bench\_right} + 2 \cdot w_{seat\_space} + w_{aisle} + w_{seat\_bench\_center}
-$
+$$
+ w_{\text{cabin}} = w_{\text{aisle}} + w_{\text{bench,left}} + w_{\text{bench,right}} + 2 \cdot w_{\text{seat space}} + w_{\text{aisle}} + w_{\text{bench,center}}
+$$
### Cabin slenderness ratio <sup>[1]</sup>
-The cabin slenderness ratio describes the ratio of cabin width to cabin length $\frac{w_{cabin}}{l_{cabin}}$.
-Whilst the cabin width is already known, the cabin length can be determined using the following equation:
-$
- l_{cabin} = \frac{n_{PAX\_per\_class}}{ab} \cdot \left[ sp + \frac{a_{service}}{w_{seat}} + \frac{a_{bulk}}{\frac{w_{aisle}}{ab} + w_{seat}} + x \cdot w_{exit} \cdot \left( \frac{ab}{n_{PAX\_per\_class}} + \frac{sp}{d_{exits}} \right) \right]
-$
+The cabin slenderness ratio describes the ratio of cabin width to cabin length and can be determined using the following equation:
+$$
+ \frac{w_{\text{cabin}}}{l_{\text{cabin}}} = \frac{n_{\text{PAX per class}}}{ab} \cdot \left[ sp + \frac{a_{\text{service}}}{w_{\text{seat}}} + \frac{a_{\text{bulk}}}{\frac{w_{\text{aisle}}}{ab} + w_{\text{seat}}} + x \cdot w_{\text{exit}} \cdot \left( \frac{ab}{n_{\text{PAX per class}}} + \frac{sp}{d_{\text{exits}}} \right) \right]
+$$
In which
+
- $x$ - factor (1 for single-aisle, 2 for wide-body)
-- $n_{PAX\_per\_class}$ - number of PAX per class
+- $n_{\text{PAX per class}}$ - number of PAX per class
- $ab$ - seat abreast
- $sp$ - seat pitch
-- $a_{service}$ - service area per PAX
-- $a_{bulk}$ - bulk area per PAX
-- $w_{exit}$ - exit width
-- $d_{exits}$ - maximum distance between two exits
+- $a_{\text{service}}$ - service area per PAX
+- $a_{\text{bulk}}$ - bulk area per PAX
+- $w_{\text{exit}}$ - exit width
+- $d_{\text{exits}}$ - maximum distance between two exits
### Cabin length
Knowing the cabin width and the cabin slenderness ratio, the cabin length (in inch) can be calculated:
-$
- l_{cabin} = \frac{w_{cabin}}{\frac{w_{cabin}}{l_{cabin}}}
-$
+$$
+ l_{\text{cabin}} = \frac{w_{\text{cabin}}}{\frac{w_{\text{cabin}}}{l_{\text{cabin}}}}
+$$
### Cabin wall thickness
The cabin wall thickness can be estimated using the following calculation:
-$
- t_{wall} = 0.02 \cdot w_{cabin} + 2.5"
-$
+$$
+ t_{\text{wall}} = 0.02 \cdot w_{\text{cabin}} + 2.5"
+$$
### Cabin floor thickness
With the use of the cabin wall thickness, the cabin floor thickness can be calculated:
-$
- t_{floor} = 1.5 \cdot t_{wall}
-$
+$$
+ t_{\text{floor}} = 1.5 \cdot t_{\text{wall}}
+$$
## Determine fuselage geometry {#fuselage-geometry}
With the calculated cabin the fuselage dimensions can be estimated.
@@ -90,85 +93,92 @@ With the calculated cabin the fuselage dimensions can be estimated.
The fuselage length can be determined via regression formulas using the cabin length (in meter).
For single-aisle aircraft:
-$
- l_{fuselage} = \frac{l_{cabin}}{0.23482756 \cdot \log l_{cabin} - 0.05106017}
-$
+$$
+ l_{\text{fuselage}} = \frac{l_{\text{cabin}}}{0.23482756 \cdot \log l_{\text{cabin}} - 0.05106017}
+$$
For wide-body aircraft:
-$
- l_{fuselage} = \frac{l_{cabin}}{0.1735 \cdot \log l_{cabin} - 0.0966}
-$
+$$
+ l_{\text{fuselage}} = \frac{l_{\text{cabin}}}{0.1735 \cdot \log l_{\text{cabin}} - 0.0966}
+$$
### Fuselage diameters
The fuselage does not necessarily have a circular cross-section. It is more common to design elliptical cross-sections. Because of that, there are several values that must be determined:
+
- Fuselage diameter in y-direction
- Fuselage diameter in negative z-direction
- Fuselage diameter in positive z-direction
#### Fuselage diameter in y-direction
-The fuselage diameter in y-direction $d_{fuselage\_y}$ can be calculated in the following way:
-$
- d_{fuselage\_y} = w_{cabin} + 2 \cdot t_{wall}
-$
+The fuselage diameter in y-direction $d_{\text{fuselage,y}}$ can be calculated in the following way:
+$$
+ d_{\text{fuselage,y}} = w_{\text{cabin}} + 2 \cdot t_{\text{wall}}
+$$
#### Fuselage diameter in negative z-direction
-The fuselage diameter in negative z-direction $d_{fuselage\_z\_neg}$ is determined by the cargo accommodation. It can be calculated in the following way.
+The fuselage diameter in negative z-direction $d_{\text{fuselage,z,neg}}$ is determined by the cargo accommodation. It can be calculated in the following way.
+
At first, the distance to the cargo bottom is calculated:
-$
- d_{to\_cargo\_bottom} = h_{max} + t_{floor} + d_{container\_to\_ceil} + o_{cabin\_floor}
-$
+$$
+ d_{\text{to cargo bottom}} = h_{\text{ULD,max}} + t_{\text{floor}} + d_{\text{container to ceil}} + o_{\text{cabin floor}}
+$$
In which
-- $h_{max}$ - maximum height of unit load device
-- $d_{container\_to\_ceil}$ - distance from the container to the ceiling
-- $o_{cabin\_floor}$ - offset cabin floor
+
+- $h_{\text{ULD,max}}$ - maximum height of unit load device
+- $t_{\text{floor}}$ - floor thickness
+- $d_{\text{container to ceil}}$ - distance from the container to the ceiling
+- $o_{\text{cabin floor}}$ - offset cabin floor
Afterwards, the distance to the lower compartment edge is estimated:
-$
- d_{to\_lower\_compartment\_edge} = d_{container\_to\_wall} + 0.5 \cdot w_{max\_at\_base}
-$
+$$
+ d_{\text{to lower compartment edge}} = d_{\text{container to wall}} + 0.5 \cdot w_{\text{base,max}}
+$$
In which
-- $d_{container\_to\_wall}$ - distance from container to wall
-- $w_{max\_at\_base}$ - maximum width at container base
+- $d_{\text{container to wall}}$ - distance from container to wall
+- $w_{\text{base,max}}$ - maximum width at container base
Based on the Pythagorean theorem, the inner fuselage diameter (that equals the hypotenuse) can be calculated:
-$
- d_{inner\_fuselage\_z\_neg} = \sqrt{(d_{to\_cargo\_bottom})^2 + (d_{to\_lower\_compartment\_edge})^2}
-$
+$$
+ d_{\text{fuselage,z,neg,inner}} = \sqrt{(d_{\text{to cargo bottom}})^2 + (d_{\text{to lower compartment edge}})^2}
+$$
Adding the wall thickness results in the fuselage diameter in negative z-direction:
-$
- d_{fuselage\_z\_neg} = d_{inner\_fuselage\_z\_neg} + t_{wall}
-$
+$$
+ d_{\text{fuselage,z,neg}} = d_{\text{fuselage,z,neg,inner}} + t_{\text{wall}}
+$$
#### Fuselage diameter in positive z-direction
-The fuselage diameter in positive z-direction $d_{fuselage\_z\_pos}$ is determined by the passenger accommodation. It can be calculated in the following way.
+The fuselage diameter in positive z-direction $d_{\text{fuselage,z,pos}}$ is determined by the passenger accommodation. It can be calculated in the following way.
+
Firstly, the inner fuselage height (equals outer cabin height) can be determined:
-$
- d_{inner\_fuselage\_z\_pos} = h_{aisle\_standing} - o_{cabin\_floor} + h_{system\_bay}
-$
+$$
+ d_{\text{fuselage,z,pos,inner}} = h_{\text{aisle,standing}} - o_{\text{cabin floor}} + h_{\text{system bay}}
+$$
In which
-- $h_{aisle\_standing}$ - passenger aisle standing height
-- $o_{cabin\_floor}$ - cabin floor offset
-- $h_{system\_bay}$ - system bay height above cabin
+
+- $h_{\text{aisle,standing}}$ - passenger aisle standing height
+- $o_{\text{cabin floor}}$ - cabin floor offset
+- $h_{\text{system bay}}$ - system bay height above cabin
Adding the wall thickness leads to the fuselage diameter in positive z-direction.
-$
- d_{fuselage\_z\_pos} = d_{inner\_fuselage\_z\_pos} + t_{wall}
-$
+$$
+ d_{\text{fuselage,z,pos}} = d_{\text{fuselage,z,pos,inner}} + t_{\text{wall}}
+$$
### Fuselage height
The total height of the fuselage can be determined by summing up the fuselage diameters in positive and negative z-direction:
-$
- h_{fuselage} = d_{fuselage\_z\_pos} + d_{fuselage\_z\_neg}
-$
+$$
+ h_{\text{fuselage}} = d_{\text{fuselage,z,pos}} + d_{\text{fuselage,z,neg}}
+$$
!!! note
If the `force_circle_cross_section` mode is selected, fuselage height and width are set to the maximum of both.
## Mass estimation {#mass-estimation}
The following masses are estimated:
+
- Fuselage structure
- Operator items
- Furnishing
@@ -176,13 +186,15 @@ The following masses are estimated:
Please refer to _Synthesis of Subsonic Airplane Design_ by E. Torenbeek<sup>[3]</sup> and the Certification Specifications<sup>[4]</sup> for further information.
!!! note
- All masses are estimated in accordance with the CPACS standard.
+ All masses are estimated in accordance with the CPACS mass standard.
<!-- ## Estimate positions and COG -->
## Generate fuselage shape {#generate-shape}
The fuselage shape is generated using the calculated data and the reference ellipses (see the [getting started](getting_started.md) page for more information). The final geometry is written to the `fuselage_design_ellipses.json` file.
+
The aircraft is divided into three sections: A cockpit section, followed by a constant section, and the tail section.
The steps of the shape generation are basically the same for all aircraft sections:
+
1. Calculate the section length as a percentage of the fuselage length<sup>*</sup>.
2. Proportionally adjust the given reference geometry to match the actual geometry using scaling factors. Therefore, separate scaling factors are calculated for
- the x-direction (lengthwise),
diff --git a/docs/documentation/sizing/fuselage_design/getting_started.md b/docs/documentation/sizing/fuselage_design/getting_started.md
index 7973557226994a3494c343be18a476f71c859f5c..ce1373f473346b76e4d452af02eba713c79e53c1 100644
--- a/docs/documentation/sizing/fuselage_design/getting_started.md
+++ b/docs/documentation/sizing/fuselage_design/getting_started.md
@@ -11,6 +11,7 @@ This section will guide you through the necessary steps to get the _fuselage\_de
It is assumed that you have the `UNICADO package` installed including the executables and UNICADO libraries.
Generally, we use two files to set or configure modules in UNICADO:
+
- The aircraft exchange file (or _acXML_) includes
- data related inputs (e.g., configuration type) and
- data related outputs (e.g., component design data).
@@ -42,6 +43,7 @@ _fuselage\_design_ can be single executed without the execution of any other mod
- ... -->
The following data should be available in the _acXML_ (2. and 3. are optional):
+
1. Requirements and specifications
- Design specification
- Configuration information
@@ -82,6 +84,7 @@ The following data should be available in the _acXML_ (2. and 3. are optional):
The _configXML_ is structured into two blocks: the control and program settings.
The control settings are standardized in UNICADO and will not be described in detail here. But to get started, you have to change at least
+
- the `aircraft_exchange_file_name` and `aircraft_exchange_file_directory` to your respective settings,
- the `console_output` at least to `mode_1`, and
- the `plot_output` to false (or define `inkscape_path` and `gnuplot_path`).
@@ -93,6 +96,10 @@ The program settings are structured like this (descriptions can be found in the
```plaintext
Program Settings
+|- Program mode
+| | - Setting use existing geometry
+| | | - Path to existing geometry_file
+| | | - Use as starting point
|- Configuration (ID="tube_and_wing")
| |- Fidelity name
| |- Method name
@@ -179,6 +186,7 @@ Program Settings
| | | | | | | | | |- Specific number of class lavatories
| | | | | | | |- Wardrobe
| | | | | | | | |- Use wardrobe for passenger class
+| | | | | | | | | | - Space per passenger
| | | | |- Passenger aisle
| | | | | |- Width
| | | | | |- Standing height
@@ -193,13 +201,16 @@ The fuselage design library contains files that are necessary to generate a vali
#### Reference ellipses
The reference aircraft ellipses are used to create the outer shape of the aircraft.
There are reference ellipses for the following sections:
+
- Cockpit section
- Constant section
- Tail sections
+
Furthermore, there is data for the reference diameter and information on scaling factors.
#### Accommodation definitions
The `accommodation_definitions.xml` file contains information on the passenger and cargo definition for the following categories:
+
- Cabin interior such as seats, galleys, trolleys, lavatories, and wardrobes as well as respective masses
- Cargo accommodation such as containers or pallets
- Emergency slides
@@ -207,6 +218,7 @@ The `accommodation_definitions.xml` file contains information on the passenger a
#### Fuselage design certification requirements
The `fuselage_design_cs_requirements.xml` file contains necessary design requirements regarding the following topics:
+
- Emergency exit definition and positioning (according to CS-25.807ff)
- Cabin design specifications such as the aisle dimensions and cross aisle overlaps (according to CS-25.807ff)
- Container arrangement
diff --git a/docs/documentation/sizing/fuselage_design/index.md b/docs/documentation/sizing/fuselage_design/index.md
index 4fee3ec3b5b95dc19a2ce9a210d09c820af11353..b470af3b9b33012e55f30d758bc82499eaf73e11 100644
--- a/docs/documentation/sizing/fuselage_design/index.md
+++ b/docs/documentation/sizing/fuselage_design/index.md
@@ -15,10 +15,12 @@ Blended-wing-body |... |... |... |under development
## A user's guide to fuselage design
The _fuselage\_design_ tool is your key to designing the aircraft's fuselage. In this user documentation, you’ll find all the information you need to understand the tool, as well as the necessary inputs and configurations to run a fuselage design from the ground up.
The following sections will walk you through the process:
+
- [Getting started](getting_started.md)
- [Run your first fuselage design](run_your_first_design.md)
For a comprehensive understanding of the tool’s functionality, the documentation is structured into two distinct sections:
+
- A [method description](design_method.md) and
- a [software architecture](software_architecture.md)
section.
diff --git a/docs/documentation/sizing/fuselage_design/run_your_first_design.md b/docs/documentation/sizing/fuselage_design/run_your_first_design.md
index 4e69db70adbb97fcaec8a962985f2a35ac160a81..f29d05baab25f3a5218c36ff66928030234794d2 100644
--- a/docs/documentation/sizing/fuselage_design/run_your_first_design.md
+++ b/docs/documentation/sizing/fuselage_design/run_your_first_design.md
@@ -3,6 +3,7 @@ Let's dive into the fun part and design a fuselage!
## Tool single execution
The tool can be executed from the console directly if all paths are set. The following will happen:
+
- [Console output](#console-output)
- [Generation of reports and plots](#reporting)
- [Writing output to aircraft exchange file](#acxml)
@@ -11,7 +12,7 @@ The tool can be executed from the console directly if all paths are set. The fol
Some of the above mentioned steps did not work? Check out the [troubleshooting](#troubleshooting) section for advices.
Also, if you need some additional information on the underlying methodology, check out the page on the [fuselage design method](design_method.md).
-So, feel free to open the terminal and run `fuselage_design.exe` to see what happens...
+So, feel free to open the terminal and run `python.exe fuselage_design.py` to see what happens...
### Console output {#console-output}
Firstly, you see output in the console window. Let's go through it step by step...
@@ -50,9 +51,10 @@ Finally, you receive information about the reports and plots created (depending
### Reporting {#reporting}
In the following, a short overview is given on the generated reports:
+
- A `fuselage_design.log` file is written within the directory of the executable
- Depending on your settings, the following output is generated and saved in the `reporting` folder, located in the directory of the aircraft exchange file:
- - an HTML report in the `report_html` folder (not implemented yet)
+ - an HTML report in the `report_html` folder
- a TeX report in the `report_tex` folder (not implemented yet)
- an XML file with additional output data in the `report_xml` folder
- plots in the `plots` folder
@@ -67,16 +69,14 @@ Aircraft exchange file
|- Component design
| |- Fuselage
| | |- Position*
-| | |- Mass properties
-| | | |- ...
+| | |- Mass properties**
| | |- Specific
| | | |- Geometry
| | | | |- Fuselage (ID="0")
| | | | | |- Name
| | | | | |- Position*
| | | | | |- Direction*
-| | | | | |- Mass properties
-| | | | | | |- ...
+| | | | | |- Mass properties**
| | | | | |- Sections
| | | | | | |- Section (ID="0")
| | | | | | | |- Name
@@ -134,7 +134,9 @@ Aircraft exchange file
| | | | | | | | | |- Payload deck required galley power
```
-<sup>*</sup> Node contains the following sub-nodes: x, y, z
+<sup>*</sup> Node has been shortened. It contains the following sub-nodes: x, y, z
+
+<sup>*</sup> Node has been shortened. It contains sub-nodes with information on the mass, inertia, and center of gravity.
### Write geometry data to .json file {#geometry-data}
The calculated geometry data is written to the `fuselage_design_ellipses.json` file and can then be used if the `use_existing_geometry` flag is set to `true`.
diff --git a/docs/documentation/sizing.md b/docs/documentation/sizing/index.md
similarity index 86%
rename from docs/documentation/sizing.md
rename to docs/documentation/sizing/index.md
index 1214ee88c1fee9c12ba20951741757a7147bbb16..949a66207a1bbfc14ab5f9d143a048d9f46d3887 100644
--- a/docs/documentation/sizing.md
+++ b/docs/documentation/sizing/index.md
@@ -15,7 +15,7 @@ The following sizing tools are available:
---
## Initial sizing
-{.overview-img align=left}
+{.overview-img align=left}
The module **initial_sizing** is used to determine a design chart regarding Top Level Aircraft Requirements and Certification Specification Requirements.
The wing-loading ($\frac{W}{S}$) and thrust to weight ratio ($\frac{T}{W}$) can be derived as the design point for further modules from the Design Chart.
Furthermore an initial estimation of the takeoff mass is done.
@@ -28,7 +28,7 @@ Furthermore an initial estimation of the takeoff mass is done.
---
## Create mission XML
-{.overview-img align=left}
+{.overview-img align=left}
The **create_mission_XML** is the third module of the UNICADO tool chain.
Its purpose is to set up the overall flight mission including e.g. a flight segment table, speed and altitude schedules, number of passengers (PAX), total payload or the engine warm up time.
For the user, possible changes in the module run configuration can be made in the related createMissionXML_conf.xml file.
@@ -42,7 +42,7 @@ The parameters comprised in this file can have different attributes as e.g. Desc
---
## Fuselage design
-{.overview-img align=left}
+{.overview-img align=left}
The **fuselage_design** module calculates characteristic parameters and generates the passenger cabin and fuselage layout for the entire aircraft project.
{.overview-item}
@@ -53,7 +53,7 @@ The **fuselage_design** module calculates characteristic parameters and generate
---
## Wing design
-{.overview-img align=left}
+{.overview-img align=left}
The **wing_design** module calculates characteristic parameter of the aircraft main wing.
{.overview-item}
@@ -64,7 +64,7 @@ The **wing_design** module calculates characteristic parameter of the aircraft m
---
## Empennage design
-{.overview-img align=left}
+{.overview-img align=left}
The **empennage_design** module calculates characteristic parameter of the empennage of the aircraft.
It takes takes the controllability as wells as the static margin of the aircraft into account and sizes the empennage accordingly.
{.overview-item}
@@ -76,7 +76,7 @@ It takes takes the controllability as wells as the static margin of the aircraft
---
## Tank design
-{.overview-img align=left}
+{.overview-img align=left}
:construction: *tbd*
{.overview-item}
@@ -88,7 +88,7 @@ It takes takes the controllability as wells as the static margin of the aircraft
---
## Propulsion design
-{.overview-img align=left}
+{.overview-img align=left}
The **propulsionDesign** module designs, integrates and analyzes the propulsion system to the aircraft.
It uses engine performance deck containing serval parameters (like thrust, fuel-flow, ...) as a function of the flight Mach number and the altitude.
The engine will be scaled by the module to match the specific thrust requirements.
@@ -105,7 +105,7 @@ Also the mass properties are analyzed.
---
## Landing gear design
-{.overview-img align=left}
+{.overview-img align=left}
The **landing_gear_design** module calculates characteristic parameters for the landing gear of entire aircraft project.
{.overview-item}
@@ -116,7 +116,7 @@ The **landing_gear_design** module calculates characteristic parameters for the
---
## Systems design
-{.overview-img align=left}
+{.overview-img align=left}
The **systems_design** is part of the tool chain in the UNICADO aircraft design environment.
It dimensions ATA chapter systems in terms of mass and energy requirement divided by hydraulic- electric- and bleed air energy requirement.
{.overview-item}
diff --git a/docs/documentation/sizing/propulsion_design/changelog.md b/docs/documentation/sizing/propulsion_design/changelog.md
index d6f5738f1468778bc0bb86eed73895825c7fc1fc..7403ff41ee9a50ae4b9369bf7c29ae5a0f94a78b 100644
--- a/docs/documentation/sizing/propulsion_design/changelog.md
+++ b/docs/documentation/sizing/propulsion_design/changelog.md
@@ -14,12 +14,11 @@ The following changes are introduced:
### Bugfixes
During the development of this release the following bugs were found and fixed:
-- When designing a *rubber* engine, the engine length was scaled incorrectly. The correct formula with ${scale_{engine}}^{0.4} $ is now implemented.
+- When designing a *rubber* engine, the engine length was scaled incorrectly. The correct formula with \f${scale_{engine}}^{0.4} \f$ is now implemented.
### Changes in the CSR designs
The implemented changes and bugfixes lead to the following changes in the results of the CSR designs.
-!!! note
- Only changes which exceed a 10 % change are listed.
+@note Only changes which exceed a 10 % change are listed.
#### CSR-02
|Parameter|Changed introduced by|Old Value|New Value|Unit|
diff --git a/docs/documentation/sizing/propulsion_design/engineering_principles.md b/docs/documentation/sizing/propulsion_design/engineering_principles.md
index e9cec0daf9a47f59d2d2e0acfdd0a3335ad84aa2..8af8f50a8cf03ea42ca0ddd57c5fc31028aae6d7 100644
--- a/docs/documentation/sizing/propulsion_design/engineering_principles.md
+++ b/docs/documentation/sizing/propulsion_design/engineering_principles.md
@@ -1,13 +1,24 @@
# Engineering principles {#engineeringprinciples}
-Designing the propulsion with this tool includes different engineering disciplines. Here a brief explanation (more information in their respective sections):
+Designing the propulsion with this tool includes different steps shown below (with more information in their respective sections):
+
+* [Engine designer](#enginedesigner): Calculates the performance of one individual engine based on the required thrust.
+* [Propulsor integrator](#propulsionintegrator): Places the engine acc. to the user's settings.
+* [Nacelle designer](#nacelledesigner): Calculates the nacelle geometry.
+* [Pylon designer](#pylondesigner): Calculates the pylon geometry.
+* [Mass analyzer](#massanalyzer): Calculates the mass properties (center of gravity, mass, and inertia) of engine, nacelle, and pylon.
+
+For these five disciplines, you can choose different methods of calculating their output. The following methods are integrated (details in the sections):
+
+| Discipline | Methods |
+|-------------------------|-------------------------------------------------------------------|
+| **Engine designer** | *Rubber* (*Empirical* and *PropulsionSystem* are in preparation) |
+| **Propulsor integrator** | *Default* |
+| **Nacelle designer** | *Default* |
+| **Pylon designer** | *Default* |
+| **Mass analyzer** | *Default* |
-- [Engine designer](#enginedesigner): calculates the performance of one individual engine based on the required thrust.
-- [Propulsor integrator](#propulsionintegrator): places the engine acc. to the user's settings.
-- [Nacelle designer](#nacelledesigner): calculates the nacelle geometry.
-- [Pylon designer](#pylondesigner): calculates the pylon geometry.
-- [Mass analyzer](#massanalyzer): calculates the mass properties (center of gravity, mass, and inertia) of engine, nacelle, and pylon.
For these five disciplines, you can choose different **methods** (or fidelities) of calculating their output. Here is an overview of the current implemented methods (details see sections):
@@ -19,66 +30,76 @@ For these five disciplines, you can choose different **methods** (or fidelities)
|Pylon designer | *Default* |
|Mass analyzer | *Default* |
-If you want to learn more about how to configure methods or generally the settings and outputs, go to the [getting started](getting_started.md).
-
-@important These disciplines are executed sequentially for EACH engine. That means that the engines are not aware of each other within the designing and analyzing. More information, see the [software architecture](software_architecture.md) section.
+If you want to learn more about how to configure methods or generally the settings and outputs, go to the [getting started](getting-started.md).
+!!! important
+ These disciplines are executed sequentially for EACH engine. That means that the engines are not aware of each other while designing and analyzing. More information, see the [software architecture](software_architecture.md) section.
## Engine designer {#enginedesigner}
+This section describes the principles of the engine designer.
### General principles {#generalprinciples}
-The **engine designer** bases its principle on the common modelling practice using
-- an _engine dataset_ (operating point **in**dependent)
-- an _engine deck_ (operating point dependent)
-- a _scale factor_
+In the engine design a dataset needs to be written into the projects directory. The following data is needed:
+
+* An engine dataset (operating point independent)
+* An engine deck (operating point dependent)
+* A scale factor
+
-The _dataset_ (also called _EngineXML_) includes parameter which are independent of the flight condition such as outer engine dimensions.
+The _dataset_ (also called _engine_xml_) includes parameter which are independent of the flight condition such as outer engine dimensions or the mass of the unscaled engine.
+
+The three-dimensional \( \text{\textit{engine deck}} \) contains engine performance data for different values of altitude \( h \), Mach number \( M_a \), and low-pressure engine spool speed \( N_1 \).
+ The most important performance parameter are thrust and fuel/energy flow. In UNICADO, the deck is split into multiple csv files. The figure shows an example with values for thrust in kilo newtons. The first block contains data for \( N_1 = 1 \) for \( M_a = 0 \ldots 0.9 \) and \( h = 0 \ldots 14000 \). The second block below is for \( N_1 = 0.95 \).
-The three-dimensional _engine deck_ contain engine performance data for different values of altitude $h$, Mach number $Ma$ and low-pressure engine spool speed $N1$. The most important performance parameter are thrust and fuel/energy flow. In UNICADO, the deck is split into multiple CSV files. The figure shows an example with values for thrust in kilo newton. The first block contains data for $N1=1$ for $Ma=0...0.9$ and $h=0...14000$. The second block below is for $N1=0.95$.
-!!! note
- Detailed information on required dataset and deck data will be updated in near future.
-The _scale factor_ is necessary because (as conceptual aircraft designer), we use the concept of a so-called _rubber engine_. That means that (depending on the method, see later) we create or assume an engine deck and provide one _scale factor_ to obtain all engine data acc. to the required thrust the engine shall provide. The figure visualized the concept:
+The _scale factor_ is necessary for the rubber method as it uses the concept of a so-called _rubber engine_. That means that (depending on the method, see later) we create or assume an engine deck and provide a _scale factor_ to scale all engine data acc. to the required thrust the engine shall provide. The figure visualized the concept:
-@attention → **As mentioned and highlighted in the figure**, there is ONE _scale factor_ **BUT** multiple exponents which differ depending on which property you want to use. E.g. for the diameter, the exponent is $0.5$ and for the mass its $0.4$. **So important to remember** that whenever you want to use engine data, you need to access it via the `engine` library. In the following, a brief explanation of the scaling concept will be given - however details are given in the library documentation.
+!!! attention
+ **As mentioned and highlighted in the figure**, there is ONE _scale factor_ **BUT** the scaling of the base values is not always linear.
+ **So important to remember** that whenever you want to use engine data, you need to access it via the `engine` library. In the following, a brief explanation of the scaling concept will be given - however details are given in the library documentation.
-So, the scaling is based on continuity principle assuming that the operating condition is constant (commonly known station numbering; assuming no pressure drop).
+The scaling is based on continuity principle assuming that the engine characteristics are constant.
$$ \textcolor{white}{T = \dot{m} \cdot (V_9 - V_0)} $$
-Therefore, thrust $T$ is proportional to the mass flow $\textcolor{white}{\dot{m}}$, which is related to the cross-sectional area $A$ of the engine.
+Therefore, thrust $T$ is proportional to the mass flow $\dot m$, which is related to the cross-sectional area $A$ of the engine.
-$$ \textcolor{white}{\dot{m}} = \rho \cdot V \cdot A = \rho \cdot V \cdot \pi \left(\frac{d}{2}\right)^2 $$
+$$ \dot m = \rho \cdot V \cdot A = \rho \cdot V \cdot \pi \frac{d}{2}^2 $$
Because area $A$ is proportional to the square of the diameter $d$ , it follows that the diameter should be proportional to the square root of the scale factor.
$$ \textcolor{white} d_{new} = d_{ref} \cdot ( \frac{T_{new}}{T_{ref}} )^{0.5} $$
-An exemplary simplified calculation (data from the V2527-A5): the current engine provides $127.27~kN$ as sea level static thrust, but for the design only $100~kN$ are needed. The scaling factor would be $0.7857$. Assuming an initial diameter $2~m$, the new diameter would be $1.773~m$ with the scaling factor of $(0.7857)^{0.5} = 0.8864$.
+An exemplary simplified calculation (data from the V2527-A5): the current engine provides $127.27~\text{kN}$ as sea level static thrust, but for the design only $100~\text{kN}$ are needed. The scaling factor would be $0.7857$. Assuming an initial diameter $2~\text{m}$, the new diameter would be $1.773~\text{m}$ with the scaling factor of $(0.7857)^{0.5} = 0.8864$.
-So, again, always access the engine data via the `engine` library to ensure that you have the correctly scaled data 🙂
-!!! note
- Actually, the sea level static thrust is not at $N1=1$ if you compare the dataset for this engine (for 110.31kN around $N1=0.95$). So the scaling factor will be slightly lower.
+The general scaling is therefore a linear scaling of the thrust. The fuel flow is scaled in the same way leading to a scaling with constant TSFC.
+
+The engine data is always accessed via the `engine` library to ensure that you have the correctly scaled data for every value. This is valid for both the non operating condition dependent variables and the values that are directly read from the deck values.
+
+!!! Note
+ Actually, the sea level static thrust is not at $N1=1$ if you compare the dataset for this engine (for $110.31~\text{kN}$ around $N1=0.95$). So the scaling factor is slightly lower.
+
### Methods description
The **engine designer** includes different methods which create/use this deck in various ways.
-- *empirical*: the initial deck is calculated based on emipirical equations
-- *rubber*: (most common approach) based on an existing deck (usually created with GasTurb), the deck is "rubberized"
-- *propulsionsystem*: with the help of the library `propulsionsystem`, different architecture can be defined and a deck created (for more information see documentation of the library)
+* *empirical*: the initial deck is calculated based on empirical equations.
+* *rubber*: (most common approach) based on an existing deck (usually created with GasTurb), the deck is "rubberized".
+* *propulsionsystem*: with the help of the library `propulsionsystem`, different architecture can be defined and a deck created (for more information see documentation of the library)
!!! note
*empirical* and *propulsionsystem* is in preparation - not implemented yet!
-For all these methods, the approach of using the _scale factor_ is the same (see explaination [here](#generalprinciples)). A deck is either first created or assumed and then data is drawn with the `engine` library with the scaling approach.
+For these methods, the approach of using the _scale factor_ is the same (see explanation [here](#generalprinciples)). A deck is either first created or an existing dataset is taken and then data is provided using the `engine` library with the scaling approach.
## Propulsion integrator {#propulsionintegrator}
Additionally to calculating the engine performance parameter, the engine has to be placed on the aircraft. The **propulsion integrator** uses the user settings from the aircraft exchange file - the following needs to be defined:
+
- parent component: wing, fuselage, empennage
- x-position (aircraft coordinate system): front or rear
- y position (aircraft coordinate system): left or right
@@ -87,6 +108,7 @@ Additionally to calculating the engine performance parameter, the engine has to
### Methods description
Here, currently only one method is implemented:
+
- *default* is based on a thesis of RWTH Aachen \cite{Ata10}
This method includes multiple empirical functions for different propulsion integration. These are the options that are currently implemented:
@@ -103,25 +125,26 @@ This method includes multiple empirical functions for different propulsion integ
For detailed information, it is referred to the thesis.
-!!! note the implementation include currently Turbofan Kerosene only
+!!! note
+ The implementation include currently Turbofan Kerosene only
## Nacelle designer {#nacelledesigner}
-After the integration, the nacelle geometry is defined (however its actually independent of the position, so the order could be changed).
+After the integration, the nacelle geometry is defined.
### Methods description
For the **nacelle designer**, only one method is implemented:
- - *default* uses the `aircraftGeometry2` library
+ - *default* uses the `aircraftGeometry2` library.
The library uses the `.dat` file defined in the _configXML_ to extrude a polygon in different sections. These sections including the origin, width, height and its profile are saved in the _acXML_. With that, every other tool can "rebuild" the geometry using the same library.
-In the current implemented method, there is no differentiation between short and long ducted nacelle. It is a polygon with 3 segments (1. and 3. segments is 25% of engine length). The diameter for the 1. and 3. segment is chosen as the maximum between fan diameter, engine width or height. The 2. segments is 25% larger.
+There is no differentiation between short and long ducted nacelle. It is a polygon with 3 segments (1. and 3. segments is 25% of engine length). The diameter for the 1. and 3. segment is chosen as the maximum between fan diameter, engine width or height. The 2. segments is 25% larger.
Keep in mind that the library defines a surface without a thickness. For more information, it is referred to the library.
-!!!note
- The implementation include currently Turbofan Kerosene only
+!!! note
+ Currently, only a kerosene turbofan engine is included.
## Pylon designer {#pylondesigner}
The pylon is the structural component to connect the engine to the aircraft.
@@ -130,18 +153,14 @@ The pylon is the structural component to connect the engine to the aircraft.
For the **pylon designer**, only one method is implemented:
- - *default* uses the `aircraftGeometry2` library
+ - *default* uses the `aircraftGeometry2` library.
-In the current method, the mounting is attached to the beginning to the nacelle to the leading edge of the wing. The length is the engine length which is extruded to the wing. the profile is, likewise for the nacelle, defined in the _configXML_.
+In the current method, the mounting is attached to the beginning to the nacelle to the leading edge of the wing. The length is the engine length which is extruded to the wing. The profile is, likewise for the nacelle, defined in the _configXML_.
-
-!!!note
- the implementation include currently Turbofan Kerosene only
-
## Mass analyzer {#massanalyzer}
-Lastly, the mass properties for engine, nacelle and pylon are calculated separate for center of gravity, mass and inertia.
+Lastly, the mass properties for engine, nacelle and pylon are calculated separately for center of gravity, mass and inertia.
### Methods description
@@ -157,6 +176,6 @@ Here, only one method is implemented:
- mass: empirical estimation
- inertia: wrt. CG with `aircraftGeometry2`lib
-!!!note
- the implementation include currently Turbofan Kerosene only
+!!! note
+ Currently, only a kerosene turbofan engine is included.
diff --git a/docs/documentation/sizing/propulsion_design/getting-started.md b/docs/documentation/sizing/propulsion_design/getting-started.md
index fafd391dbf79ade980793db140d0bcfc7be33527..7ee24ca9599ed8f3385162011424a7c02909265d 100644
--- a/docs/documentation/sizing/propulsion_design/getting-started.md
+++ b/docs/documentation/sizing/propulsion_design/getting-started.md
@@ -17,7 +17,6 @@ It is assumed that you have the `UNICADO Package` installed including the execut
3. Open terminal and run **propulsion_design**
Following will happen:
-
- you see output in the console window
- a HTML report is created in the directory of `aircraft_exchange_file_directory` (no plots of engine if they are turned off)
- results are saved in the `/aircraft_exchange_file/component_design/propulsion`
@@ -29,7 +28,7 @@ Following will happen:
Generally, we use 2 files to set or configure in UNICADO:
- the aircraft exchange file (or _acXML_) includes
- - data related inputs (e.g. thrust, offtakes or type of engine)
+ - data related inputs (e.g. thrust-to-weight ratio, MTOM, average bleed and shaft offtakes or type of engine)
- data related outputs (e.g. engine position)
- the configuration file `propulsion_design_conf.xml` (also _configXML_) includes
- control settings (e.g. enable/disable generating plots)
@@ -42,11 +41,11 @@ Generally, we use 2 files to set or configure in UNICADO:
**Inputs**:
Following is needed from the _acXML_:
-1) the total required thrust,
-2) the system off-takes,
-3) the user settings of the propulsion architecture
+1) the total required thrust using the thrust-to-weight ratio and MTOM,
+2) the average system off-takes for the bucket-curve,
+3) the user settings of the propulsion architecture.
-Naturally, the propulsion need an assumption for thrust or power to be designed. Currently, in UNICADO, the requirement is set via the tool _initialSizing_. Here, initial estimation based on the TLARs are calculated like the thrust-to-weight via an design chart or the maximum take-off mass based on regressions. For this, **propulsion_design** currently assumes:
+The propulsion design tool is based on the overall thrust or power the propulsion needs to be designed for. The thrust_share input divides the overall thrust to the single propulsors. In the first run of the UNICADO workflow, the tool _initialSizing_ estimates the thrust-to-weight-ratio for this. Afterwards, the tool _constraint_analysis_ updates the thrust to weight ratio by calculation the performance values using actual aircraft data. This assures the total thrust is sufficient to certification boundary conditions. With the thrust-to-weight ratio, which is calculated for the sea level static thrust, the propulsors are designed.
The sea level static thrust $T_0$ is given by:
@@ -61,7 +60,7 @@ Where:
!!! note
This might change with new propulsion architectures!
-Not only the thrust is important, but also the system off-takes. Current engine provide power to the systems and therefore, the thrust specific consumption can increase. To include that, the nodes `average_bleed_air_demand` and `average_bleed_air_demand` in `/aircraft_exchange_file/component_design/systems/specific/`are necessary (is set to default values if not existing).
+The most important parameter is the thrust-to-weight-ratio. Another input are the average system off-takes. Current engined provide power to different systems and therefore, the thrust specific consumption will increase. To include that, the nodes `average_bleed_air_demand` and `average_bleed_air_demand` in `/aircraft_exchange_file/component_design/systems/specific/` are read (is set to default values if not existing).
Additionally, the user settings need to be defined. In the node `/aircraft_exchange_file/requirements_and_specifications/design_specification`, both `energy_carriers` and `propulsion` need to be filled out (for more information on the variables, please read the description in the _acXML_).
@@ -82,8 +81,8 @@ Propulsion
| |- Energy Carrier ID
| |- Thrust Share
```
-Let's assume you want to design an aircraft with 5 engine - 2 on each side of the wing and one in the empennage. Additionally, you want to use 3 energy carrier: hydrogen, kerosene and battery-electric.
-For that, you need to define 3 energy carriers with each a type and a density with $ID=[0,1,2]$. Then you create 5 propulsor nodes with $ID=[0,...,4]$ and assign them each an a powertrain, type, ..., and thrust share. E.g. Engine 0 shall be a kerosene-powered turbofan in the empennage with a thrust share of $10\%$. Then it has the position with `parent_component=empennage`, `x=front`, `y=mid`, `z=in`. If the type of the energy carrier with ID=0 is set to kerosene, you need to assign `energy_carrier_id=0`. Also `powertrain=turbo`, `type=fan`, and `thrust_share=0.1`. Then Engine 1 could be a hydrogen-powered turboprop located under the left front inner wing with a thrust share of $25\%$. Then it has the position with `parent_component=wing`, `x=front`, `y=left`, `z=under`. If the type of the energy carrier with ID=1 is set to hydrogen, you need to assign `energy_carrier_id=1`. Also `powertrain=turbo`, `type=prop`, and `thrust_share=0.25`. The same procedure needs to be done for the other 3 engine.
+Let's assume you want to design an aircraft with 5 engine - 2 on each side of the wing and one in the empennage. Additionally, you want to use 3 energy carriers: hydrogen, kerosene and battery-electric.
+For that, you need to define 3 energy carriers with each a type and a density with $ID=[0,1,2]$. Then you create 5 propulsor nodes with $ID=[0,...,4]$ and assign them each a powertrain, type, ..., and thrust share. E.g. Engine 0 shall be a kerosene-powered turbofan in the empennage with a thrust share of $10\%$. Then it has the position with `parent_component=empennage`, `x=front`, `y=mid`, `z=in`. If the type of the energy carrier with ID=0 is set to kerosene, you need to assign `energy_carrier_id=0`. Also `powertrain=turbo`, `type=fan`, and `thrust_share=0.1`. Then Engine 1 could be a hydrogen-powered turboprop located under the left front inner wing with a thrust share of $25\%$. Then it has the position with `parent_component=wing`, `x=front`, `y=left`, `z=under`. If the type of the energy carrier with ID=1 is set to hydrogen, you need to assign `energy_carrier_id=1`. Also `powertrain=turbo`, `type=prop`, and `thrust_share=0.25`. The same procedure needs to be done for the other 3 engines.
**Outputs**: The results are saved in the _acXML_ node `/aircraft_exchange_file/component_design/propulsion`.
@@ -154,7 +153,7 @@ Program Settings
| | |- Profile
| |- Integration
```
-You can choose the method for each discipline, the path for your engine data base, and different technology factors. To be highlighted, is the `Propulsion ID=Default` node. This is a default for all engines defined in the _acXML_ (see next paragraph). E.g. if you define 3 engines for an aircraft, both will use the same assumptions in the default setting. In case you want that the 3. engine is been calculated with e.g. another method, you can create a new `propulsion` node and assign the same `ID` value as set for the _acXML_ `ID`.
+You can choose the method for each discipline, the path for your engine data base, and different technology factors. To be highlighted, is the `Propulsion ID=Default` node. This is a default for all engines defined in the _acXML_ (see next paragraph). E.g. if you define 3 engines for an aircraft, all will use the same assumptions in the default setting. In case you want that the 3. engine is been calculated with e.g. another method, you can create a new `propulsion` node and assign the same `ID` value as set for the _acXML_ `ID`.
## Minimal required aircraft exchange file input {#acXML}
diff --git a/docs/documentation/sizing/propulsion_design/index.md b/docs/documentation/sizing/propulsion_design/index.md
index 9caaff81ee51c1e3563458a952533ae99313abc9..a568bb88aa9b89f6baa867e570e399c4ac6b4638 100644
--- a/docs/documentation/sizing/propulsion_design/index.md
+++ b/docs/documentation/sizing/propulsion_design/index.md
@@ -1,14 +1,13 @@
# Introduction {#mainpage}
-The tool _propulsion_design_ is one of the core design tools in UNICADO. The overall goal is the design the propulsion based on...
-
+The tool _propulsion_design_ is one of the core design tools in UNICADO. The overall goal is the design the propulsion system based on...
- the architecture (e.g. 2 turbofan at rear fuselage, 4 fuel cell prop engine over the front wing) set by the user and,
-- the total required thrust and system off-takes.
-This tool is exciting!🔥 because the propulsion is THE critical component providing the thrust or power, enabling to propel the aircraft forward and move through the skies.ðŸŒ
+- the total required thrust and system off-takes calculated within the aircraft design loop.
+This tool is exciting!🔥 because the propulsion is one of the critical components in the aircraft design loop. It provides the thrust or power, enabling powered flight of the aircraft letting it move through the skies.ðŸŒ
-To give you a general taste, here are a few illustrations of possible concepts.
+There are different propulsion architectures for the aircraft conceptual design process. To give you a general taste, here are a few illustrations of possible concepts.
-The [getting started](getting_started.md) gives you a first insight in how to execute the tool and how it generally works. To understand how the tools works in detail, the documentation is split into a [engineering principles](engineering_principles.md) and a [software architecture](software_architecture.md) section.
+The [getting started](getting_started.md) gives you a first insight in how to execute the tool and how it generally works. To understand how the tools works in more detail, the documentation is split into a [engineering principles](engineering_principles.md) and a [software architecture](software_architecture.md) section.
Prior to that, let's summarize what the tool can currently do and what is planned (terms like _method_ or _strategy_ will be explained in the sections):
@@ -19,7 +18,12 @@ Prior to that, let's summarize what the tool can currently do and what is planne
|kerosene-powered turboprop | |strategy integrated, but methods missing |
|hydrogen-powered turboprop | |strategy integrated, but methods missing |
-(*) order: engine designer/ nacelle designer/ pylon designer/ propulsion integrator/ mass analyzer
+Order:
-So let's get started!
+ 1. engine designer
+ 2. nacelle designer
+ 3. pylon designer
+ 4. propulsion integrator
+ 5. mass analyzer
+So let's get started!
diff --git a/docs/documentation/sizing/propulsion_design/software_architecture.md b/docs/documentation/sizing/propulsion_design/software_architecture.md
index 294c85b487ec2fd2b1c5f3815f92876669f0db95..7995d60cc6570e0d69f3e9877e3f4b25f74bff5c 100644
--- a/docs/documentation/sizing/propulsion_design/software_architecture.md
+++ b/docs/documentation/sizing/propulsion_design/software_architecture.md
@@ -2,7 +2,7 @@
## Software Architecture Overview
-The software architecture is structured into various modules and packages, each handling specific task. Below is a description of the main components (some classes, interfaces etc. are left out to keep it understandable for now - for more information see the [class diagram](figures/class_diagram.png) or the source code):
+The software architecture is structured into various modules and packages, each handling specific task. Below is a description of the main components (some classes, interfaces etc. are left out to keep it understandable for now - for more information see the [class diagram](figure/class_diagram.png) or the source code):
- classes:
- **propulsionDesign** is like the "coordinator" responsible for the overall propulsion system design including _initialize_, _run_, _update_, _report_ and _save_ (inherits from `Module` class from **moduleBasics**). These include e.g. method selection function for each disciplines
@@ -21,9 +21,9 @@ The software architecture is structured into various modules and packages, each
Some additional words on the **propulsionStrategy**:
-As you might also see in the [class diagram](figures/class_diagram.png), the core of it is the functor `operator()` for specific engine types to allow the `engine` object to be used as functions. This object is, depending on the user settings, based on the propulsion type classes (e.g. `Turbofan<Kerosene>`). As also shown in @ref propulsion.md, the type is combined with 3 "building blocks"
+As you might also see in the [class diagram](figure/class_diagram.png), the core of it is the function `operator()` for specific engine types to allow the `engine` object to be used as functions. This object is, depending on the user settings, based on the propulsion type classes (e.g. `Turbofan<Kerosene>`). As also shown in @ref propulsion.md, the type is combined with 3 "building blocks"
+
- *powertrain*: Way the power is generated from the source: turbo, electric, fuel_cell
- -
- *type*: Type of main thrust generator: fan or prop
- *energy_carrier*: kerosene, liquid_hydrogen, battery (handled over IDs)
diff --git a/docs/documentation/sizing/systems_design/getting-started.md b/docs/documentation/sizing/systems_design/getting-started.md
index 6099db298cf73b1470539d83d4b3e27e28345dff..204d27b4867761580f557d63ff2e019d741f0c2e 100644
--- a/docs/documentation/sizing/systems_design/getting-started.md
+++ b/docs/documentation/sizing/systems_design/getting-started.md
@@ -34,7 +34,7 @@ Three input files are required for **systems_design**:
- tank
- propulsion
- number of flight and cabin crew
-- the configuration file `initial_sizing_conf.xml` (or _configXML_) includes
+- the configuration file `systems_design_conf.xml` (or _configXML_) includes
- control settings (e.g. enable/disable generating plots)
- program settings (e.g. define system architecture, set parameters for individual systems)
- the mission file (`design_mission.xml`, `study_mission.xml` or `requirement_mission.xml`) is required since **systems_design** calculates the required system power for each mission step.
diff --git a/docs/documentation/sizing/tank_design/getting_started.md b/docs/documentation/sizing/tank_design/getting_started.md
index 4b81a6aaf1d26f550149beb30bea9891fbabb0b4..7933ecf90658d15be8f465d7782c28ab73fd2164 100644
--- a/docs/documentation/sizing/tank_design/getting_started.md
+++ b/docs/documentation/sizing/tank_design/getting_started.md
@@ -12,9 +12,11 @@ This section will guide you through the necessary steps to get the _tank\_design
It is assumed that you have the `UNICADO package` installed including the executables and UNICADO libraries.
Generally, we use two files to set or configure modules in UNICADO:
+
- The aircraft exchange file (or _acXML_) includes
- data related inputs (e.g., required energy, component design data) and
- data related outputs (e.g., tank positions).
+
- The module configuration file `tank_design_conf.xml` (also _configXML_) includes
- control settings (e.g., enable/disable generating plots) and
- program settings (e.g., information on buffers).
@@ -37,12 +39,14 @@ Thus, it must be ensured that this data is available. More information on requir
## Aircraft exchange file requirements {#aircraft-exchange-file}
To single execute the _tank\_design_ module, we need an _acXML_ file that already contains the output data from the following tools:
+
- _wing\_design_
- _empennage\_design_
- _fuselage\_design_
-- _mission\_analysis_ - _tank\_design_ execution also possible without mission analysis data (an assumption is made to calculate mission energy amount)
+- _mission\_analysis_ <sup>*</sup>
The following data should then be available in the _acXML_:
+
1. Requirements and specifications
- Design specification
- Configuration information: Configuration type, tank definition ([see below](#configuring-tank-design-parameters-in-the-aircraft-exchange-file))
@@ -59,11 +63,26 @@ The following data should then be available in the _acXML_:
!!! note
When the UNICADO workflow is executed the tool is run automatically. In this case, all the required data should be available anyway.
+<sup>*</sup> The _tank\_design_ execution is also possible without mission analysis data. Alternatively, the following assumption is used to calculate the mission fuel amount:
+
+$$
+ m_{\text{fuel}} = n_{\text{PAX}} \cdot R \cdot \frac{E}{100 \text{ km}}
+$$
+
+In which
+
+- $n_{\text{PAX}}$ - number of passengers
+- $R$ - range in km
+- $E$ - energy demand (3.35 liter per PAX per 100 km)
+
+Using the volumetric energy density of kerosene, the initial energy demand can then be calculated.
+
## Configuring tank design parameters in the aircraft exchange file {#configuring-tank-design-parameters-in-the-aircraft-exchange-file}
The desired tank configuration is defined by the user in the aircraft exchange file. The information can be found in the `aircraft_exchange_file/requirements_and_specifications/design_specification/configuration/tank_definition` block.
### The ID `tank_element`
Each tank is configured in the _acXML_ as one element (ID element) with the following parameters:
+
- `energy_carrier_ID`: ID of energy carrier to obtain which fuel is to be stored in the tanks.
- `location`: Aircraft component where the tank is located (valid options depend on the energy carrier).
- `position`: Position at the desired location (valid options depend on location and energy carrier).
@@ -91,6 +110,7 @@ For aircraft configurations with a kinked wing, the "wing tank configuration" co
| 7 | Fuselage | Center | Also referred to as 'additional center tank'. |
For example, to define a valid combination for the wing center tank, set the following parameters for the ID element `ID="0"`:
+
- `energy_carrier_ID` to `0` (it is assumed that `0` equals kerosene)
- `location` to `wing`
- `position` to `center`
@@ -110,6 +130,7 @@ The following table provides an overview on possible tank configurations. As can
| Wing | Fuselage - Center | - | wing_with_additional_center_tank |
| Wing | Horizontal stabilizer - Total | Fuselage - Center | wing_with_additional_center_and_trim_tank|
| Wing | Fuselage - Center | Horizontal stabilizer - Total | wing_with_additional_center_and_trim_tank|
+
Note: "Wing" always refers to either the combinations 1 to 5 of the table in the previous section ("wing tank configuration").
#### Possible tank configurations: Aircraft with singe trapezoidal wing
@@ -122,6 +143,7 @@ The following table provides an overview on possible tank configurations. As can
| Wing | Fuselage - Center | - | wing_with_additional_center_tank |
| Wing | Horizontal stabilizer - Total | Fuselage - Center | wing_with_additional_center_and_trim_tank|
| Wing | Fuselage - Center | Horizontal stabilizer - Total | wing_with_additional_center_and_trim_tank|
+
Note: "Wing" always refers to the combinations 1, 2, and 4 of the table in the previous section ("wing tank configuration").
#### Example: Minimum tank configuration
@@ -242,6 +264,7 @@ tbd. :construction:
The _configXML_ is structured into two blocks: the control and program settings.
The control settings are standardized in UNICADO and will not be described in detail here. But to get started, you have to change at least
+
- the `aircraft_exchange_file_name` and `aircraft_exchange_file_directory` to your respective settings,
- the `console_output` at least to `mode_1`, and
- the `plot_output` to false (or define `inkscape_path` and `gnuplot_path`).
diff --git a/docs/documentation/sizing/tank_design/index.md b/docs/documentation/sizing/tank_design/index.md
index 9b314d9f92d723a6505e1d1fc798ad7513a8d6fb..a8af86c9cad9fa9612e475df211cf2a522549205 100644
--- a/docs/documentation/sizing/tank_design/index.md
+++ b/docs/documentation/sizing/tank_design/index.md
@@ -13,10 +13,12 @@ Blended-wing-body |... |... |... |under development
## A user's guide to tank design
The _tank\_design_ tool is your key to designing the aircraft's fuel storage. In this user documentation, you’ll find all the information you need to understand the tool, as well as the necessary inputs and configurations to run a tank design from the ground up.
The following sections will walk you through the process:
+
- [Getting started](getting_started.md)
- [Run your first tank design](run_your_first_tank_design.md)
For a comprehensive understanding of the tool’s functionality, the documentation is structured into two distinct sections:
+
- A [method description](tank_design_method.md) and
- a [software architecture](software_architecture.md)
section.
diff --git a/docs/documentation/sizing/tank_design/run_your_first_tank_design.md b/docs/documentation/sizing/tank_design/run_your_first_tank_design.md
index 400fe0aff55f1fca7fb2fdf15af407988070d875..ee39580f507ee760f6e503da9c24aab5d4c660d7 100644
--- a/docs/documentation/sizing/tank_design/run_your_first_tank_design.md
+++ b/docs/documentation/sizing/tank_design/run_your_first_tank_design.md
@@ -3,6 +3,7 @@ Let's dive into the fun part and design some tanks!
## Tool single execution
The tool can be executed from the console directly if all paths are set. The following will happen:
+
- [Console output](#console-output)
- [Generation of reports and plots](#reporting)
- [Writing output to aircraft exchange file](#acxml)
@@ -10,7 +11,7 @@ The tool can be executed from the console directly if all paths are set. The fol
Some of the above mentioned steps did not work? Check out the [troubleshooting](#troubleshooting) section for advices.
Also, if you need some additional information on the underlying methodology, check out the page on the [tank design method](tank_design_method.md).
-So, feel free to open the terminal and run `tank_design.exe` to see what happens...
+So, feel free to open the terminal and run `python.exe tank_design.py` to see what happens...
### Console output {#console-output}
Firstly, you see output in the console window. Let's go through it step by step...
@@ -46,11 +47,11 @@ The tool continues with the calculation of the wing tank entities - in this exam
2024-12-10 13:05:45,847 - PRINT - Energy check: Wing center tank necessary to store required energy amount.
2024-12-10 13:05:45,847 - PRINT - Energy check: Energy demand covered.
```
-After the wing tank design there is an energy check to review whether the required mission energy can be stored in the tanks. If the energy demand would not be covered up until this point, an energy check would be occur after the calculation of every subsequent tank.
+After the wing tank design there is an energy check to review whether the required mission energy can be stored in the tanks. If the energy demand would not be covered up until this point, an energy check would be carried out after the calculation of every subsequent tank.
```
2024-12-10 13:05:45,848 - PRINT - Additional center tank design started...
-2024-12-10 13:05:45,848 - PRINT - Additional center tank (tank_5) calculated. Volume (energy) available: 3,068.50 L (99,189.26 MJ)
+2024-12-10 13:05:45,848 - PRINT - Additional center tank (tank_5) calculated. Volume (energy) available: 3,068.50 L (103,818.10 MJ)
2024-12-10 13:05:45,849 - PRINT - Additional center tank design completed.
2024-12-10 13:05:45,849 - PRINT - Additional center tank is generated but unnecessary to store required energy amount.
```
@@ -78,15 +79,14 @@ Finally, you receive information about the reports and plots created (depending
### Reporting {#reporting}
In the following, a short overview is given on the generated reports:
+
- A `tank_design.log` file is written within the directory of the executable
- Depending on your settings, the following output is generated and saved in the `reporting` folder, located in the directory of the aircraft exchange file:
- - an HTML report in the `report_html` folder (not implemented yet)
+ - an HTML report in the `report_html` folder
- a TeX report in the `report_tex` folder (not implemented yet)
- - an XML file with additional output data in the `report_xml` folder (not written since no more data output necessary)
+ - an XML file with additional output data in the `report_xml` folder (currently, only a rough output file is generated with the routing information but without any additional data)
- plots in the `plots` folder (not implemented yet)
-@warning Steffi: Check if additional output written
-
### Write data to the aircraft exchange file {#acxml}
!!! note
The _acXML_ is an exchange file - we agreed on that only data will be saved as output that is needed by another tool!
@@ -96,33 +96,22 @@ Results are saved in the aircraft exchange file at the `/aircraft_exchange_file/
Aircraft exchange file
|- Component design
| |- Tank
-| | |- Position
-| | | |- x
-| | | |- y
-| | | |- z
-| | |- Mass properties
-| | | |- ...
+| | |- Position*
+| | |- Mass properties**
| | |- Specific
| | | |- Additional fuselage length
| | | |- Tank (ID="0")
| | | | |- Name
| | | | |- Designator
-| | | | |- Position
-| | | | | |- x
-| | | | | |- y
-| | | | | |- z
-| | | | |- Direction
-| | | | |- Mass properties
-| | | | | |- ...
+| | | | |- Position*
+| | | | |- Direction*
+| | | | |- Mass properties**
| | | | |- Maximum energy capacity
-| | | | |- Energy required for mission energy
+| | | | |- Energy capacity required for mission
| | | | |- Geometry
| | | | | |- Cross section (ID="0")
| | | | | | |- Name
-| | | | | | |- Position
-| | | | | | | |- x
-| | | | | | | |- y
-| | | | | | | |- z
+| | | | | | |- Position*
| | | | | | |- Shape
| | | | | | |- Height
| | | | | | |- Width
@@ -133,5 +122,9 @@ Aircraft exchange file
| | | | |- ...
```
+<sup>*</sup> Node has been shortened. It contains the following sub-nodes: x, y, z
+
+<sup>*</sup> Node has been shortened. It contains sub-nodes with information on the mass, inertia, and center of gravity.
+
## Troubleshooting {#troubleshooting}
- The tool does not run properly? *Make sure you have all the paths set up correctly and the specified elements exist.*
diff --git a/docs/documentation/sizing/tank_design/tank_design_method.md b/docs/documentation/sizing/tank_design/tank_design_method.md
index dd7c72b4e9c8b65f7740fba5b8aad1e40fa6343c..d5ea8d7d8dd5c2f284c45af6398500b287be7757 100644
--- a/docs/documentation/sizing/tank_design/tank_design_method.md
+++ b/docs/documentation/sizing/tank_design/tank_design_method.md
@@ -1,5 +1,6 @@
# Calculation method
The task of the _tank_design_ module differs slightly depending on the energy carrier:
+
- [Kerosene](#kerosene-tanks) - Determine the maximum fuel capacity of the aircraft using its geometry.
- [Liquid hydrogen](#liquid-hydrogen-tanks) - Size tanks to ensure that the required amount of fuel is available.
@@ -42,6 +43,15 @@ The obelisk method simplifies the wing by dividing it into several volumes. Depe
+##### Obelisk geometry
+
+The geometry of the obelisks is obtained based on the wing geometry.
+Knowing the chord length $l_\text{chord}$ and the thickness-to-chord ratio, the maximum profile thickness $h_\text{max}$ can be obtained.
+The actual thickness $h_1$ is calculated using the a-to-d factor (user input).
+The width of the obelisk $w_1$ is defined as the distance between the front $p_\text{fs}$ and the rear spar $p_\text{rs}$ of the wing.
+
+
+
The obelisk volume can be determined using two different approaches that are described in the following.
The user can select the desired method via the following node in the `program_settings` section of the _confXML_:
`configuration[@ID="tube_and_wing"]/specific/kerosene_tank_design_parameter/obelisk_calculation_method`.
@@ -53,10 +63,13 @@ The user can select the desired method via the following node in the `program_se
The volume can be calculated using the following equation:
-$
+
+$$
V_{\text{obelisk}} = \frac{l}{3} \cdot \left( S_1 + S_2 + \frac{h_1 \cdot w_2 + h_2 \cdot w_1}{2}\right)
-$
+$$
+
In which
+
- $l$ - length
- $S_1$, $S_2$ - end face areas
- $h_1$, $w_1$ - height and width of end face $S_1$
@@ -91,9 +104,10 @@ The Simpson's rule is a method of numerical integration that is often used to ca
cross-sections are known at different positions. In the case of an obelisk - i.e. a body with square or rectangular
cross-sections that vary along the height - the volume is integrated as the sum of the cross-sectional areas $S(x)$
along the length $l$:
-$
+
+$$
V_{\text{obelisk}} = \int_0^l S(x) \, dx
-$
+$$
If the cross-sectional areas $S(x)$ at $i+1$ uniformly distributed points are known (which is the case for the
tank design), Simpson's rule can be applied.
@@ -104,35 +118,39 @@ interpolation (see following figure). Each tank is thus divided into two section
The tank volume can therefore be determined using a simplified Simpson's rule:
-$
+
+$$
V_{\text{obelisk}} = \frac{l}{6.0} \cdot (S_{1} + 4.0 \cdot S_{12} + S_{2})
-$
+$$
#### Calculate net tank volume {#net-tank-volume}
The volume must then be converted from cubic meter to liter. A portion of the volume of the obelisk is lost to the internal structure of the integral tanks (e.g., ribs), with a reduction factor $ f_{\text{volume,usable}} = 0.95$.
Additionally, the expansion of the fuel due to heating must be considered, with a temperature expansion allowance of $ a_{\text{temperature,expansion}} = 0.95$. Thus, the wing tank volume is calculated as:
-$
+
+$$
V_{\text{tank}} = f_{\text{volume,usable}} \cdot a_{\text{temperature,expansion}} \cdot V_{\text{obelisk}}
-$
+$$
!!! note
As the wing has a vent tank at each wing tip to allow for the thermodynamic expansion of the fuel, this factor is `1.0` for the wing tanks.
#### Calculate energy {#calculate-energy}
Using the volumetric energy density of kerosene $\eta_{\text{v,kerosene}}$, the energy contained in each tank can be determined:
-$
+
+$$
V_{\text{tank}} = \eta_{\text{v,kerosene}} \cdot V_{\text{obelisk}}
-$
+$$
### Additional center tank
The module allows the installation of an additional center tank in the form of an LD3-45 container. The process includes the following steps:
+
1. **Height check**: The program first verifies whether the cargo compartment has sufficient height to accommodate the container.
- - If insufficient height is detected: All output values related to the center tank are set to zero.
- - If sufficient height is detected: The installation proceeds.
+ - If insufficient height is detected: All output values related to the center tank are set to zero.
+ - If sufficient height is detected: The installation proceeds.
2. **Installation placement**: The LD3-45 container is positioned 10 cm behind the end of the landing gear bay, aligning approximately with the trailing edge of the wing.
3. **Container data**: The volume and dimensions of the LD3-45 container are predefined and referenced from the Lufthansa Cargo website<sup>[2]</sup>.
-The energy contained in an additional center tank is calculated by first determining the usable volume and then taking into account the volumetric energy density of kerosene. With the known factors $ f_{\text{volume,usable}}$ and $a_{\text{temperature,expansion}}$, the usable volume of an additional center tank results in $V_{\text{ACT,usable}} = 3068.5\text{ L}$ which is equal to an energy amount of $ E_{\text{ACT}} = 99189262.5\text{ MJ}$.
+The energy contained in an additional center tank is calculated by first determining the usable volume and then taking into account the volumetric energy density of kerosene. With the known factors $ f_{\text{volume,usable}}$ and $a_{\text{temperature,expansion}}$, the usable volume of an additional center tank results in $V_{\text{ACT,usable}} = 3,068.5\text{ L}$ which is equal to an energy amount of $ E_{\text{ACT}} = 103,818.1\text{ MJ}$.
#### Limitations
**Single Center Tank Limit:** The program currently supports the calculation of only one additional center tank. Attempts to add more tanks will not be processed.
diff --git a/docs/get-involved/build-environment/windows.md b/docs/get-involved/build-environment/windows.md
index e6ad8940172254a4a34604279ebedc507ec3a8d1..2b8c24fc60055c55a7231d7d5f42dea174d402bc 100644
--- a/docs/get-involved/build-environment/windows.md
+++ b/docs/get-involved/build-environment/windows.md
@@ -31,7 +31,7 @@ The tools used are:
!!! warning
Please install version **3.11.8**, select all the default options, and check the option to add Python to *PATH* & make sure to include the debug binaries!
-
+
### Python Dependencies
The only Python dependency we have is `pipenv`, which is used to manage the Python environment.
@@ -69,7 +69,7 @@ Afterwards, you can freely decide on installing nice extensions such as:
- Download the build tools from Microsoft: [Download Build Tools ](https://visualstudio.microsoft.com/downloads/?q=build+tools#build-tools-for-visual-studio-2022){:target="_blank"}
The page should look something like this:
-
+
- Execute the installer and install at least these components:
- *Desktop development with C++*
diff --git a/docs/get-involved/build/general.md b/docs/get-involved/build/general.md
index 11d0a9de4952f4b4e7c2c940f7e8f3ea35d06dfc..1d6d6a58548a6fead12a171bb240851e13318c1a 100644
--- a/docs/get-involved/build/general.md
+++ b/docs/get-involved/build/general.md
@@ -36,7 +36,7 @@ Since **CMake** is independent of the used platform, it needs to figure out on w
**CMake** also comes with a **GUI** which can be used for the configuration process.
The GUI looks like this:
-
+
It works the same as the command line interface.
You have to specify the path to the source files and to the build directory.
@@ -63,4 +63,4 @@ The other options shown in the GUI are specific to **CMake**.
Once the configuration is complete and build files are generated, the build step compiles the modules into executables or libraries. This is where CMake hands over control to the underlying compiler, which translates the C++ and Python components of UNICADO into a runnable format.
The different modules are called *targets* in the **CMake** language.
-After this step, you’ll have compiled executables and/or libraries ready to run or integrate into the UNICADO workflow. :simple-cmake: :heart:
\ No newline at end of file
+After this step, you’ll have compiled executables and/or libraries ready to run or integrate into the UNICADO workflow. :simple-cmake: :heart:
diff --git a/docs/get-involved/contribute.md b/docs/get-involved/contribute.md
index 8ec14b9abc6e852aa21c9f30dd5242b84a4e450b..b09b59007f38bb88fae950695f457d3d21f04048 100644
--- a/docs/get-involved/contribute.md
+++ b/docs/get-involved/contribute.md
@@ -20,10 +20,11 @@ This is how you can actually make a difference:
The flowchart below illustrates the Merge Request workflow, along with the commands used at each stage.
<figure>
- <img src="../../assets/images/merge_request_workflow.png" alt="merge-request" width="500" style="border: 2px solid black;">
+ <img src="site:assets/images/merge_request_workflow.png" alt="merge-request" style="width: 80%; height: auto;" >
<figcaption>Merge request workflow</figcaption>
</figure>
+
You cloned/forked the UNICADO Package successfully acc. to [Get Source Code](get-source-code.md). Nice! You want to make a change, e.g. fixing a bug or creating a new feature, so you create a *issue* (see also [types of contribution](#contributions)). Then you :point_up: create a feature branch, change the code and create a merge request (here a [how to](merge-request.md)). An automatic CI/CD pipeline is triggered, which helps your selected reviewer to make sure that request is ok. If it is accepted and ready-to-land :airplane:, the documentation is automatically updated. Nicely done :+1:
## Types of contribution {#contributions}
diff --git a/docs/get-involved/style/cpp-modularization.md b/docs/get-involved/modularization/cpp-modularization.md
similarity index 100%
rename from docs/get-involved/style/cpp-modularization.md
rename to docs/get-involved/modularization/cpp-modularization.md
diff --git a/docs/get-involved/style/python-modularization.md b/docs/get-involved/modularization/python-modularization.md
similarity index 94%
rename from docs/get-involved/style/python-modularization.md
rename to docs/get-involved/modularization/python-modularization.md
index d8a39ba66424cbfa72f285a1bf97fd8af4cc8b0a..d9851f66fcee031521fb878421310b185bcd3ccf 100644
--- a/docs/get-involved/style/python-modularization.md
+++ b/docs/get-involved/modularization/python-modularization.md
@@ -138,7 +138,7 @@ This means, for example:
# Code modularity (Python-only modules) {#code-modularity-python-only-modules}
In the following, the modularized structure of a Python module is explained using the `cost_estimation` module. The according folder structure is shown in the following picture. It is also available for download.
-
+
!!! warning
Check, if images displayed correctly here!
@@ -154,7 +154,7 @@ The following **layers** are selected for cost calculation:
3. Calculation method (e.g., `operating_cost_estimation_tu_berlin`, green folder)
4. Energy carrier (e.g., `kerosene` or `liquid_hydrogen`, grey folder) - **USER LAYER** (This is where the magic happens! :dizzy:)
-
+
!!! warning
Check, if images displayed correctly here!
@@ -201,7 +201,7 @@ rAircraftDesign
## Files that require changes by the module manager
The code is designed to be highly generalized, meaning that only a few files need changes by the module manager. These files are shown in the following image and are discussed below in more detail. In some parts of the code, dynamic import commands and function names are generated, with examples provided at relevant points to illustrate how these commands work.
-
+
!!! warning
Check, if images displayed correctly here!
@@ -211,8 +211,8 @@ The code is designed to be highly generalized, meaning that only a few files nee
- Customize the module configuration file name
- Adjust the `runtime_output_string`
-
-
+
+
!!! warning
Check, if images displayed correctly here!
@@ -221,25 +221,25 @@ The code is designed to be highly generalized, meaning that only a few files nee
- Update the layer description in the docString
- Customize the layer description within `layer_description_dict`. If a layer is unknown (e.g., `user_layer`), set it to 'None' rather than a path and call the relevant function (e.g., `read_energy_carrier`) as indicated (see lines 69 and following).
-
-
+
+
!!! warning
Check, if images displayed correctly here!
-**Example for `module_import_name`**
+**Example for `module_import_name`**
In this example, `module_import_name` at line 68 would be: `src.tube_and_wing.empirical.operating_cost_estimation_tu_berlin`.
-**Example for the import command**
-To import a module from `usermethoddatapreparation.py` at line 74, the command is as follows:
+**Example for the import command**
+To import a module from `usermethoddatapreparation.py` at line 74, the command is as follows:
`src.tube_and_wing.empirical.operating_cost_estimation_tu_berlin.usermethoddatapreparation`.
### The `data_postprocessing` (`datapostprocessing.py`)
- Modify `paths_to_key_parameters_list`
- Adjust `module_key_parameters_dict`
-
-
+
+
!!! warning
Check, if images displayed correctly here!
@@ -258,7 +258,7 @@ Users are free to structure the code within these files but must ensure that all
More detailed instructions for required changes are available within the docStrings of each corresponding file.
-
+
!!! warning
Check, if images displayed correctly here!
@@ -275,7 +275,7 @@ The Python framework in this project has a customized logging function, which bu
| `runtime_output.error` | 40 | For serious issues where the code can still continue | `runtime_output.error("Error: Add some text here.")` |
| `runtime_output.critical` | 50 | For critical issues that terminate the code (exit code 1) | `runtime_output.critical("Error: Add some text here.")` |
-Instead of using Python's built-in `print` function, use these logging options to ensure all outputs are appropriately documented in the log file according to user settings.
+Instead of using Python's built-in `print` function, use these logging options to ensure all outputs are appropriately documented in the log file according to user settings.
## Logging configuration in the module configuration file
User settings for logging behavior can be configured in the module configuration file under `console_output/value` and `log_file_output/value`. The available modes are as follows:
@@ -303,7 +303,7 @@ The necessary steps are listed below. Please ensure to read the respective expla
- **Windows:** `python -m pip install --upgrade pip`
2. Navigate to the `AircraftDesign/unicado_python_library` folder (illustrated below) to set up the required folder structure.
-
+
!!! warning
Check, if images displayed correctly here!
@@ -312,7 +312,7 @@ The necessary steps are listed below. Please ensure to read the respective expla
In `unicado_python_library`, create a new subfolder for the package. Follow this naming convention:
- **Format:** `py[name of package]package` (all lowercase, without underscores)
- **Example:** `pymodulepackage`
-
+
Then, navigate into this subfolder.
## Step 2: Create a `pyproject.toml` file
@@ -348,7 +348,7 @@ Inside the package folder, create a `src` subfolder to hold the `.py` files (mod
- **Convention:** Each `.py` file should correspond to a single module, named in this format:
- **Format:** `[module name]module.py` (all lowercase, no underscores)
- **Example:** `datapreprocessingmodule.py`
-
+
Modules can contain several functions. Once files are set up, return to the main package folder before proceeding.
## Step 6: Execute installation command
@@ -372,4 +372,4 @@ The modules should now be ready to use. You can import the functions from the m
`from datapostprocessingmodule import paths_and_names`
# Testing with Python {#testing-with-python}
-tbd. :construction:
\ No newline at end of file
+tbd. :construction:
diff --git a/docs/partners.md b/docs/partners.md
index 11c9299d28254135864bded2fa8e1caee287ea6c..8037bb19c686f65e4ee173b382b4bab1b383cb78 100644
--- a/docs/partners.md
+++ b/docs/partners.md
@@ -7,7 +7,7 @@
<!-- RWTH Aachen -->
<div class="grid-item card" markdown="1">
<p align="center">
- <a href="https://www.rwth-aachen.de/"><img src="../assets/images/logos/RWTH.svg" alt="Logo RWTH" width="150"/></a>
+ <a href="https://www.rwth-aachen.de/"><img src="site:assets/images/logos/RWTH.svg" alt="Logo RWTH" width="150"/></a>
</p>
---
@@ -17,7 +17,7 @@
<!-- TU Berlin -->
<div class="grid-item card" markdown="1">
<p align="center">
- <a href="https://www.tu.berlin"><img src="../assets/images/logos/TUB.svg" alt="Logo TUB" width="150"/></a>
+ <a href="https://www.tu.berlin"><img src="site:assets/images/logos/TUB.svg" alt="Logo TUB" width="150"/></a>
</p>
---
@@ -28,7 +28,7 @@
<!-- TU Braunschweig -->
<div class="grid-item card" markdown="1">
<p align="center">
- <a href="https://www.tu-braunschweig.de"><img src="../assets/images/logos/TUBS.svg" alt="Logo TUBS" width="150"/></a>
+ <a href="https://www.tu-braunschweig.de"><img src="site:assets/images/logos/TUBS.svg" alt="Logo TUBS" width="150"/></a>
</p>
---
@@ -38,7 +38,7 @@
<!-- TUHH -->
<div class="grid-item card" markdown="1">
<p align="center">
- <a href="https://www.tuhh.de"><img src="../assets/images/logos/TUHH.svg" alt="Logo TUHH" width="150"/></a>
+ <a href="https://www.tuhh.de"><img src="site:assets/images/logos/TUHH.svg" alt="Logo TUHH" width="150"/></a>
</p>
---
@@ -48,7 +48,7 @@
<!-- TUM -->
<div class="grid-item card" markdown="1">
<p align="center">
- <a href="https://www.tum.de"><img src="../assets/images/logos/TUM.svg" alt="Logo TUM" width="150"/></a>
+ <a href="https://www.tum.de"><img src="site:assets/images/logos/TUM.svg" alt="Logo TUM" width="150"/></a>
</p>
---
@@ -58,7 +58,7 @@
<!-- USTUTT -->
<div class="grid-item card" markdown="1">
<p align="center">
- <a href="https://www.uni-stuttgart.de"><img src="../assets/images/logos/USTUTT.svg" alt="Logo USTUTT" width="150"/></a>
+ <a href="https://www.uni-stuttgart.de"><img src="site:assets/images/logos/USTUTT.svg" alt="Logo USTUTT" width="150"/></a>
</p>
---
@@ -71,7 +71,7 @@
<div class="grid-container" markdown="1">
<div class="grid-item card" markdown="1">
<p align="center">
- <a href="https://www.tuwien.at"><img src="../assets/images/logos/TUW.png" alt="Logo TUW" width="150"/></a>
+ <a href="https://www.tuwien.at"><img src="site:assets/images/logos/TUW.png" alt="Logo TUW" width="150"/></a>
</p>
---
diff --git a/docs/tutorials/changing-design-specifications.md b/docs/tutorials/changing-design-specifications.md
index 80987335511da343fe80868fb8ffeaf09ab7f0dc..b630282511bdc72c145363b60c7b2c802ea0ff25 100644
--- a/docs/tutorials/changing-design-specifications.md
+++ b/docs/tutorials/changing-design-specifications.md
@@ -8,8 +8,8 @@ date: 2024-12-04
When running the UNICADO workflow, the aircraft is sized based on requirements and specifications defined in the aircraft exchange file. This file holds initial parameters in an XML format, which ensure the final aircraft meets the design objectives and that can be modified to accommodate alternative goals or mission profiles.
## Overview of the Aircraft Exchange File
-In its hierarchy, the most important parameters for the definition of the aircraft can be found in the sections
-`requirements_and_specifications/design_specifications`, where requirements regarding the configuration, transport task, propulsion and technologies are defined, and in `requirements_and_specifications/requirements`, where top-level aircraft requirements (TLARs) are defined.
+In its hierarchy, the most important parameters for the definition of the aircraft can be found in the sections
+`requirements_and_specifications/design_specifications`, where requirements regarding the configuration, transport task, propulsion and technologies are defined, and in `requirements_and_specifications/requirements`, where top-level aircraft requirements (TLARs) are defined.
```
<aircraft_exchange_file>
└── <requirements_and_specifications>
@@ -28,10 +28,10 @@ When modifying parameters, it is important to consider the intended design objec
├── <lower_boundary> Lower boundary, can not be changed
└── <upper_boundary> Upper boundary, can not be changed
```
-When changing parameters, it is recommended to change only one parameter in the beginning to study how a change will affect the sizing of the aircraft.
+When changing parameters, it is recommended to change only one parameter in the beginning to study how a change will affect the sizing of the aircraft.
## Most important parameters
-The following tables showcase the most important parameters that affect the sizing of the aircraft, as well as their location within the aircraft exchange file structure. For additional details about other parameters, refer to the descriptions provided within their corresponding XML tags.
+The following tables showcase the most important parameters that affect the sizing of the aircraft, as well as their location within the aircraft exchange file structure. For additional details about other parameters, refer to the descriptions provided within their corresponding XML tags.
Definition of design mission related parameters:
@@ -49,7 +49,7 @@ Definition of transport task related parameters:
|Additional Cargo Mass|Mass of cargo which does not belong to passengers. Make sure, that the maximum structural payload mass is changed accordingly.|`requirements_and_specifications/design_specification/transport_task/cargo_definition/additional_cargo_mass`|
|Maximum structural payload mass|Maximum structual payload mass which can be carried by the aircraft. Must at least include passenger and cargo mass.|`requirements_and_specifications/requirements/top_level_aircraft_requirements/maximum_structrual_payload_mass`|
-Definition of the aircraft configuration:
+Definition of the aircraft configuration:
|Parameter|Description|Path in Aircraft Exchange File|
|---|---|---|
@@ -84,9 +84,9 @@ Following assumptions are considered:
`delta_ISA` is therefore a mission related parameters which changes the starting point of ISA - commonly adapted over the temperature. The figure shows it examplary for ISA+10.
<figure markdown>
- {width="500"}
+ {width="500"}
<figcaption>Altitude over temperature for ISA and ISA+10</figcaption>
</figure>
---
-<sup>[1]</sup> International Organization for Standardization, Standard Atmosphere, ISO 2533:1975, 1975.<br>
\ No newline at end of file
+<sup>[1]</sup> International Organization for Standardization, Standard Atmosphere, ISO 2533:1975, 1975.<br>
diff --git a/docs/tutorials/standalone.md b/docs/tutorials/standalone.md
index 8ccc96134806cae0387abb82b182fab4f2ff8918..916fbe650063935d31ab9dbac7ad0ac12b81ac7e 100644
--- a/docs/tutorials/standalone.md
+++ b/docs/tutorials/standalone.md
@@ -1 +1,2 @@
-@todo here should be a video showing the standalone workflow
\ No newline at end of file
+!!! tip "ToDo"
+ here should be a video showing the standalone workflow
diff --git a/docs/workflow.md b/docs/workflow.md
index 5052deaae07e783bb0ca0fb36fd663efa03ffef1..e93cb60dcdcc8005f9bafa461e1a9aa58e211dfb 100644
--- a/docs/workflow.md
+++ b/docs/workflow.md
@@ -15,8 +15,8 @@ To open the UNICADO workflow, launch RCE and with `File → Open Projects from F
<figcaption>UNICADO Workflow</figcaption>
</figure>
-The workflow can be executed with `Run → Execute Workflow...`, after which the user is asked to name the current run and specify the installation path to the installed python version on the computer. Once defined, the workflow will beginn execution. Due to the various iteration steps involved in the aircraft design process, the execution may take some time to complete.
-During execution, log entries of the different modules can be seen in the Workflow Console in RCE. When finished, the results can be found in `workflowResults` in the installation directory of UNICADO, where they can be viewed and analyzed.
+The workflow can be executed with `Run → Execute Workflow...`, after which the user is asked to name the current run and specify the installation path to the installed python version on the computer. Once defined, the workflow will beginn execution. Due to the various iteration steps involved in the aircraft design process, the execution may take some time to complete.
+During execution, log entries of the different modules can be seen in the Workflow Console in RCE. When finished, the results can be found in `workflowResults` in the installation directory of UNICADO, where they can be viewed and analyzed.
## Configuration Settings
The workflow in RCE can be executed to perform different tasks, depending on the user set `program_settings/design_case_settings/design_mode` value in the configuration file, which can be found in `workingDirectoryRCE/UNICADOworkflow/unicado_workflow_conf.xml`. The following table gives an overview over the different design modes currently possible to run:
@@ -35,4 +35,4 @@ To speed up the execution of the workflow, the configuration file also holds fur
|Parameter|Description|Path in Workflow Configuration File
|---|---|---|
|Convergence Criteria|Max. allowed relative change to the last iteration to achieve convergence|`program_settings/design_case_settings/iteration_settings/convergence_criteria`
-|Max. number of Iterations|Max. allowed number of iterations before the workflow is aborted|`program_settings/design_case_settings/iteration_settings/max_number_of_iterations_before_exit`
\ No newline at end of file
+|Max. number of Iterations|Max. allowed number of iterations before the workflow is aborted|`program_settings/design_case_settings/iteration_settings/max_number_of_iterations_before_exit`
diff --git a/mkdocs.yml b/mkdocs.yml
index 0282aaec330d6d90ecb0e90ef9c5af27319e20dc..e41bb5fa4e6b96493b34fc5277fd7b91130c93be 100644
--- a/mkdocs.yml
+++ b/mkdocs.yml
@@ -18,6 +18,7 @@
site_name: UNICADO # The name of the site, displayed in the header.
repo_url: https://git.rwth-aachen.de/unicado/unicado-package # Link to the Git repository, will appear in the header.
repo_name: UNICADO Repository # Name for the Git repository link in the header.
+site_url: "https://unicado.pages.rwth-aachen.de/unicado.gitlab.io/" # The actual site url -> IMPORTANT: site-urls relies on this (site: will be replaced directly)!
# === Site configuration ===
markdown_extensions:
@@ -47,12 +48,16 @@ markdown_extensions:
- pymdownx.keys # Adds special styling for keyboard key indicators.
- pymdownx.mark # Adds highlighting functionality for text.
- pymdownx.tilde # Enables strikethrough formatting.
+ - pymdownx.arithmatex:
+ generic: true
# Additional JavaScript files to include for rendering mathematical notation
extra_javascript:
- assets/javascripts/katex.js # Local KaTeX script.
- https://cdnjs.cloudflare.com/ajax/libs/KaTeX/0.16.7/katex.min.js # CDN KaTeX script (same as local but hosted externally).
- https://cdnjs.cloudflare.com/ajax/libs/KaTeX/0.16.7/contrib/auto-render.min.js # KaTeX auto-render script (converts Latex syntax in formatted math).
+ - assets/javascripts/mathjax.js # Local MathJax script
+ - https://unpkg.com/mathjax@3/es5/tex-mml-chtml.js # MathJax renderer can be used for more complex formulas
# Additional CSS files to include for styling of website and mathematical notations (font, size etc.)
extra_css:
@@ -62,6 +67,7 @@ extra_css:
# === Plugins ===
plugins:
- search
+ - site-urls
- mkdoxy:
projects:
propulsion_design:
@@ -144,6 +150,14 @@ plugins:
FILE_PATTERNS: "*.cpp *.h"
RECURSIVE: True
EXTRACT_ALL: YES
+ aerodynamic_analysis:
+ src-dirs: ../aircraft-design/aerodynamic_analysis/
+ full-doc: true
+ output: docs/aerodynamic_analysis
+ doxy-cfg:
+ FILE_PATTERNS: "*.cpp *.h"
+ RECURSIVE: True
+ EXTRACT_ALL: YES
aircraftGeometry2:
src-dirs: ../aircraft-design/libs/aircraftGeometry2/
full-doc: true
@@ -152,6 +166,14 @@ plugins:
FILE_PATTERNS: "*.cpp *.h"
RECURSIVE: True
EXTRACT_ALL: YES
+ engine:
+ src-dirs: ../aircraft-design/libs/engine/
+ full-doc: true
+ output: docs/engine
+ doxy-cfg:
+ FILE_PATTERNS: "*.cpp *.h"
+ RECURSIVE: True
+ EXTRACT_ALL: YES
- glightbox # Plugin for lightbox-style image and content viewing.
@@ -171,7 +193,11 @@ theme:
features:
- navigation.instant
- navigation.top
+ - navigation.path
- navigation.tabs
+ - navigation.tabs.sticky
+ - navigation.indexes
+ - toc.follow
# Additional links (social) to display in the header
extra:
@@ -185,7 +211,7 @@ theme:
nav: # Customizes the main navigation structure of the site.
- Home: index.md # Main page of the site.
- Download: # Top-level navigation item for "Download".
- - Getting Started: download/getting_started.md # Link to the getting started page.
+ - Getting Started: download/getting-started.md # Link to the getting started page.
- Installation: download/installation.md # Link to the installation page.
- Cleared for Take-Off: download/takeoff.md # Link to the takeoff/getting started page.
- Tutorials:
@@ -196,7 +222,7 @@ nav: # Customizes the main navigation struc
- Overview: documentation/overview.md # Overview of modules.
- Aircraft Design:
- Sizing:
- - Modules: documentation/sizing.md # Link to aircraft sizing documentation.
+ - documentation/sizing/index.md # Link to aircraft sizing documentation.
- Initial Sizing:
- Introduction: documentation/sizing/initial_sizing/index.md
- Getting Started: documentation/sizing/initial_sizing/getting-started.md
@@ -207,6 +233,10 @@ nav: # Customizes the main navigation struc
- initial_sizing/namespaces.md
- initial_sizing/files.md
- initial_sizing/functions.md
+ - Create Mission XML:
+ - Introduction: documentation/sizing/create_mission_xml/index.md
+ - Getting Started: documentation/sizing/create_mission_xml/getting_started.md
+ - Mission Steps: documentation/sizing/create_mission_xml/mission_steps.md
- Fuselage Design:
- Introduction: documentation/sizing/fuselage_design/index.md
- Getting Started: documentation/sizing/fuselage_design/getting_started.md
@@ -265,7 +295,7 @@ nav: # Customizes the main navigation struc
- Design Method: documentation/sizing/landing_gear_design/design_method.md
- Run your First Design: documentation/sizing/landing_gear_design/run_your_first_design.md
- Software Architecture: documentation/sizing/landing_gear_design/software_architecture.md
- # # - API Reference: # TODO define for Python
+ # - API Reference: # TODO define for Python
- Systems Design:
- Introduction: documentation/sizing/systems_design/index.md
- Getting Started: documentation/sizing/systems_design/getting-started.md
@@ -276,8 +306,18 @@ nav: # Customizes the main navigation struc
- systems_design/namespaces.md
- systems_design/files.md
- systems_design/functions.md
- - Analysis:
+ - Analysis:
- Modules: documentation/analysis.md # Link to analysis module page.
+ - Mission Analysis:
+ - Introduction: documentation/analysis/mission_analysis/index.md
+ - Getting Started: documentation/analysis/mission_analysis/getting_started.md
+ - Mission Methods: documentation/analysis/mission_analysis/methods.md
+ - Mission Steps: documentation/analysis/mission_analysis/mission_steps.md
+ - API Reference:
+ - mission_analysis/classes.md
+ - mission_analysis/namespaces.md
+ - mission_analysis/files.md
+ - mission_analysis/functions.md
- Weight and Balance Analysis:
- Introduction: documentation/analysis/weight_and_balance_analysis/index.md
- Basic Concepts: documentation/analysis/weight_and_balance_analysis/basic-concepts.md
@@ -301,8 +341,15 @@ nav: # Customizes the main navigation struc
- ecological_assessment/namespaces.md
- ecological_assessment/files.md
- ecological_assessment/functions.md
+ - Aerodynamic Analysis:
+ - Introduction: documentation/analysis/aerodynamic_analysis/getting_started.md
+ - Aerodynamic Principles: documentation/analysis/aerodynamic_analysis/aerodynamic_principles.md
+ - Software Architecture: documentation/analysis/aerodynamic_analysis/software_architecture.md
+ - Constraint Analysis:
+ - Introduction: documentation/analysis/constraint_analysis/index.md
+ - Principles: documentation/analysis/constraint_analysis/principles.md
- Libraries:
- - Overview: documentation/libraries.md # Link to libraries overview.
+ - documentation/libraries/index.md # Link to libraries overview.
- AircraftGeometry2:
- Introduction: documentation/libraries/aircraftGeometry2/index.md
- Getting Started: documentation/libraries/aircraftGeometry2/getting-started.md
@@ -312,6 +359,8 @@ nav: # Customizes the main navigation struc
- aircraftGeometry2/namespaces.md
- aircraftGeometry2/files.md
- aircraftGeometry2/functions.md
+ - engine:
+ - Introduction: documentation/libraries/engine/index.md
- Utilities: documentation/additional_software.md
- Workflow: 'workflow.md' # Link to the workflow page.
- Get Involved:
@@ -330,8 +379,8 @@ nav: # Customizes the main navigation struc
- Include Libraries: get-involved/including-libraries.md
- CMake Presets: get-involved/cmake-presets.md
- Module Development:
- - Module Structure in c++: get-involved/style/cpp-modularization.md
- - Module Structure in Python: get-involved/style/python-modularization.md
+ - Module Structure in c++: get-involved/modularization/cpp-modularization.md
+ - Module Structure in Python: get-involved/modularization/python-modularization.md
- Style Guide:
- C++: get-involved/style/cpp.md
- Python: get-involved/style/python.md