systems.md



Implemented Aircraft System Models
These system models are relevant if you choose to use the standard strategy of systems design.

ATA 21: Environmental Control System
The environmental control system model implemented is powered by electric power and bleed air from the engines.
Methods
The power required by the environmental control system is calculated based on the heat loads of all systems, the heat from the sun and from the passengers. The ECO-Mode allows to reduce the required bleed air by 25%. Air conditioning is switched off during takeoff.
The mass of the ECS depends on the bleed air mass flow in the design point. Calculation method from LTH and Howe. The mass is broken down into the components ducts, air conditioning pack, outlet, ram inlet, vents, and misc. according to factors determined by Koeppen¹.
The CoG of the ECS is determined with the assumption that its located in the belly fairing.
Required Input Parameters

Airflow per PAX [kg/s]
Recirculation [-]: percentage of cabin air that is reused (0.0 - 1.0)
Heat Convection [W/(m^2*K)]: heat convection over aircraft skin (based on Airbus air conditioning system design)
Cabin Temperature [K]
Specific Heat Flow from Sun [W/m^2]
Window Area [m^2]: Area of a single window
Heat per PAX [W]: heat emitted per person
Heat per Light Length [W/m]
Efficiency Factor of the Air Conditioning Pack [-]
Heat Capacity Air [J/(kg*K)]
Off Take Off: Switch to turn of ACP during take off
ECO Mode: Switch for ECO Mode reduces bleed air requirement by 25%


ATA 24: Electric System
Methods
CheckUserInput() checks if generator sources exist.
The required power of the electric system is the power lost through inefficiencies. The efficiency factor for the electric system and for the generators are considered.
The mass calculation method from Steinke is based on the maximum required electric power and the cable length. The cable length is defined as 2*fuselage length (main bus from front to back, two bus systems) + connection of the engines to the avionics bay. A factor is applied to the mass depending on the design range of the aircraft (short or long range).
Required Input Parameters

Efficiency Factor [-]
Maximum Relative Power [-]: ratio of maximum permanent power and maximum required power of all generators
Specific Cable Mass [kg/m]
Number of Electric Circuits [-]
Number of Generators [-]
For each Generator:

name
type (IDG, APUG, ...)
source type (hydraulic, engine, APU)
source ID
efficiency
operation factor (share of total power in normal operation)


ATA 25: Furnishing System
Furnishing system includes the power required for the galleys and the inflight entertainment system (IFE). The mass of all furnishing is calculated in fuselage design and read from the aircraft exchange file.
Methods
Power calculation is done based on user inputs.
!!! important
Because the mass of furnishing items is already calculated in fuselage_design, it is not changed in systems_design and there is no scaling factor available in systems_design to adapt this mass. If you wish to make adaptions, you can do this in fuselage_design.
Required Input Parameters

Galley Load Fraction during Takeoff [-]:
Electric Load Analysis for A320 suggest a value of 0.2¹


Galley Load Fraction during Cruise [-]:
Electric Load Analysis for A320 suggest a value of 0.7¹

Galley Load Fraction during Descent [-]:
Electric Load Analysis for A320 suggest a value of 0.2¹

Galley Location [m]
Non Personal IFE Power [W]: General power for IFE
Personal IFE Power [W]: Power for IFE per PAX
Personal IFE Load Fraction Climb [-]:
Electric Load Analysis for A320 suggest a value of 0.58¹

Personal IFE Load Fraction Cruise [-]:
Electric Load Analysis for A320 suggest a value of 1¹

Personal IFE Load Fraction Descent [-]:
Electric Load Analysis for A320 suggest a value of 0.5¹


ATA 26: Fire Protectrion System
Does not require power!
Methods
Mass calculation is based on propulsion type and MTOM (Torenbeek Tab. 8-12).
Required Input Parameters
None.

ATA 27: Flight Control System
The flight control system is modeled in great detail down to the individual actuators of the control surfaces. The calculation is devided into segments according to the control surfaces:

Ailerons
Spoilers
Elevators
Rudders
Trimmable Horizontal Stabilizers (THSAs)
Flaps
Slats

The actuator architecture is defined by the user in the configuration file but the control surface geometry is read from the aircraft exchange file. The control surface geometry from the aircraft exchange file and the actuator architecture are checked against each other to make sure they match.
Methods
The function setHorizontalContrSurfGeo is used to read the geometry of horizontal control surfaces (e.g. ailerons, elevator) while the function setRudderTemp is used for the rudder. For the THSA there is a specialized function to read the geometry: setTHSA_AsControlSurface. These functions are used to calculate the area moment about the hinge line of each control surface and to determine the reference points of the actuators. This requires the geometry of the aerodynamic surface the control surfaces is mounted on and the geometry of the control surface itself, as well as the maximum deflection, deflection speed and number of actuators on the control surface.
The design power of primary flight control surfaces is calculated based on the control surface geometry and hinge moment, the wing loading, and the maximum operating speed of the aircraft. The hinge moment is determined according to [NASA76] (equations on page 26).
For the high lift system there's a separate class calculating the mass and power required (class highLiftSystem). This class and all other classes used for the calculation of the high lift system are in the folder src/aircraftSystems/highLiftSystem. The high lift devices are sorted into trailing edge and leading edge panels. For each device the actuation moments for the actuators, their work load and power are calculated. The mass consists of the masses of the actuation and support, the PCU, the wing tip brakes, the gear boxes, the torque shaft, and the torque limiters. The power consists of the actuation power and PCU power.
There are two functions for the mission power allowing for detailed power calculation at each mission step or the calculation of the average mission power. Power calculation is the same as for the design power but with the flight speed at the given mission step.
The mass of the primary flight control system is based on [NASA76, p. 42-48].
Required Input Parameters
There are some general inputs and then inputs for each control surface type. The general inputs are:

Switch whether loads for each mission step are calculated (otherwise average values)
Switch for electrical flight control system
Common installation weight factor [-]
Default actuator:

Power Source (type and ID)
Operation Mode (active, standby or damping during normal operation)
Efficiency [-]
Standby Power [W]
Active Power [W]


Specific inputs for each control surface type:

Weight Factor for Electric Flight Control System [-]
(This factor will only be applied if the switch for electric flight control system is set true!)
Switch if default actuator architecture should be used
Default Actuator Architecture:

Number of Actuators per Control Surface [-]
Default Deflection Speed [deg/s]


Manual Acturator Archtitecture:

Number of Control Surfaces [-]
For each Control Surface:

Side (left or right)
Deflection Speed [deg/s]
Actuator Layout:

Number of Actuators [-]
For each actuator same parameters as for default actuator


Acceptable control surface names are:

ailern / droop_aileron
spoiler_air / spoiler_ground
flap / slotted_flap / ADHF / fowler / double_fowler / triple_fowler / special / morphing_trailing_edge
slat / krueger / droop_nose / morphing_droop_nose / special

Note: not all of these devices are implemented. If they aren't implemented a default will be used and a warning issued.

ATA 28: Fuel System
The fuel system does not contain the tank mass!
Methods
The power of the electric powered pumps of the fuel system is calculated according to Buente, based on the MTOM.
The mass of the fuel system is the mean value of the results from the Raymer and Torenbeek method. A technology factor of 0.7 is applied to the Raymer method.
The CoG is assumed to be the same as that of the wing.
Required Input Parameters
None

ATA 29: Hydraulic System
Methods
CheckUserInput() checks if pump sources exist.
The required power of the hydraulic system is the power lost through inefficiencies. The efficiency factor for the hydraulic system and for the pumps are considered.
The mass calculation method from Steinke¹ is based on the maximum required hydraulic power, the OME, the fluid mass and the ducting length. The ducting length is composed of the lengths from the engines to the belly fairing, the front gear to the back gear, 2 * the length of the wing trailing edge, 2 * the trailing edges of the horizontal and vertical tail plane. The total length is then doubled to account for the backflow.
Required Input Parameters

Pressure [Pa]
Efficiency Factor [-]
Relative Maximum Power [-]: Ratio of maximum permanent power and maximum required power of all pumps
Specific Ducting Mass [kg/m]
Specific Pump Mass [kg/W]
Number of Hydraulic Circuits [-]
For each circuit:

Compartment Reference Point [m]
Number of Pumps [-]
For each pump:

Name
Type (electric driven, enginge driven, RAT)
Pump Efficiency [-]
Operation Factor [-]: Percentage of total pump power in normal operation
Power Source (type and ID)


ATA 30: Ice and Rain Protection System
There are two implemented models for the ice and rain protection system - a conventional one powered by bleed air and an electric one.

Conventional ATA 30
Methods
Mass is calculated as a percentage of the OME defined by the user.
Anti-Icing is typically applied from the kink of the wing and ends at the outer most leading edge device. If there are no leading edge devices the anti-icing is applied up until the wing tip. The system is design for the continuous maximum icing condition. The design altitude and design mach number are used to determine if there are icing conditions. The calculation of the liquid water content of the air is based on a method from CS-25 Appendix C (Figure 1). This method is valid between -30°C and 0°C. If the altitude's temperature lies outside of these boundaries (e.g. due to setting a delta temperature to the international standard atmosphere (ISA) in the acXML) the temperature is set to -30° or 0°, respectively. This ensures that the bleed air required by the anti-icing system is considered in the design loop. If there are icing conditions the external heat flux is calculated based on the water catch, the external heat transfer coefficient, the vapour pressure and from that the skin temperature. The required bleed air mass flow can then be calculated with the external heat flux, the inner skin temperature, the bleed air temperature, and the bleed air efficiency.
Required Input Parameters

Switch to turn off anti-icing
Top Operating Altitude [m]
Engine Anti Ice Bleed Air Mass Flow [kg/s]
Skin Thickness [m]
Relative Half Span Width to Start Anti-Icing [-]
Heat Conductivity Wing [W/(m*K)]
Drop Diameter [micro meter]
Mass Percentage of OME [-]


Electric ATA 30
Mass is calculated as a percentage of the OME defined by the user.
Methods
Same calculation for the external heat flux. From there the required electric power can be calculated with the electro thermic efficiency and the user defined electric power.
Required Input Parameters
Same as for conventional +

Electric Power Consumption Departure [W]:
Electric Load Analysis for A320 suggest a value of 14026.9 W¹

Electric Power Consumption Cruise [W]:
Electric Load Analysis for A320 suggest a value of 13070.9 W¹

Electric Power Consumption Approach [W]:
Electric Load Analysis for A320 suggest a value of 14026.9 W¹

Electric Power Consumption Land [W]
Electric Load Analysis for A320 suggest a value of 7192.9 W¹

Efficiency Electro Thermic Anti-Icing [-]


ATA 32: Landing Gear System