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SLRE is a Python package that provides simulation tools for the preliminary design of liquid rocket engines.

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Features

Features include: - Integrated CEA analysis through ceapy.py - Integrated thermodynamic property querying through CoolProp - Generation of nozzle geometry (conical or Rao bell nozzle approximation) - Generation of combustion chamber geometry - One-dimensional regenerative cooling analysis - Engine cycle analysis

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Installation

I have largely halted development of SLRE, but may revisit it in the future. The code is functional, but I have not extensively tested for bugs and usability, so be aware of this if you decide to use it.

To install:

1. Clone the repository:

shell git clone https://github.com/TomAveyard/SLRE.git

1. Create a python virtual environment. If using VS Code, follow the following instructions: https://code.visualstudio.com/docs/python/environments

2. Install the requirements using:

shell pip install -r requirements.txt

If using VS Code you may have already been prompted to do this when setting up the virtual environment

1. You can test that everything is running correctly by running testScript.py in the main folder. This should result in an engine being solved for and the following output:

```

Engine Performance

Specific Impulse: 217.25 s Exit Velocity: 2131.23 m/s Thrust Coefficient: 1.39

C Star: 1552.26 m/s

Mass Flow Rates

Fuel Mass Flow Rate: 1.04 kg/s Oxidiser Mass Flow Rate: 3.65 kg/s

Total Mass Flow Rate: 4.69 kg/s

Temperatures

Chamber Temperature: 2812.47 K Throat Temperature: 2545.13 K

Exit Temperature: 1579.48 K

Sizes

Chamber Radius: 68.1 mm Chamber Area: 14570.48 mm^2 Throat Radius: 34.05 mm Throat Area: 3642.62 mm^2 Exit Radius: 62.11 mm

Exit Area: 12118.89 mm^2

Power Balance

Regenerative Cooling Heat Power: 1267.47 kW Fuel Pump Power: -9.86 kW Total Pump Power: -9.86 kW Turbine Power: 10.21 kW

Power Balance: 0.36 kW

Fuel Line Component Outputs

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Fuel Pipe 1 Results

Substance: propanol Mass Flow Rate: 1.04 kg/s Inlet Temperature: 298.0 K Inlet Pressure: 3.0 Bar Outlet Temperature: 298.0 K Outlet Pressure: 2.91 Bar

Pressure Loss: 0.09 Bar

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Fuel Pump Results

Substance: propanol Mass Flow Rate: 1.04 kg/s Inlet Temperature: 298.0 K Inlet Pressure: 2.91 Bar Outlet Temperature: 300.27 K Outlet Pressure: 40.0 Bar Temperature Rise: -2.27 K Pressure Rise: 37.09 Bar Head Rise: 482.3 m

Power: -9.86 kW

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Fuel Pipe 2 Results

Substance: propanol Mass Flow Rate: 1.04 kg/s Inlet Temperature: 300.27 K Inlet Pressure: 40.0 Bar Outlet Temperature: 300.27 K Outlet Pressure: 39.93 Bar

Pressure Loss: 0.07 Bar

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Fuel Regen Cooling Results

Substance: propanol Mass Flow Rate: 1.04 kg/s Inlet Temperature: 300.27 K Inlet Pressure: 39.93 Bar Outlet Temperature: 581.9 K Outlet Pressure: 39.28 Bar Temperature Rise: 281.64 K Pressure Loss: 0.65 Bar Max Heat Flux: 8103.84 kW/m^2

Max Gas Side Wall Temp: 660.98 K

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Fuel Pipe 3 Results

Substance: propanol Mass Flow Rate: 1.04 kg/s Inlet Temperature: 581.9 K Inlet Pressure: 39.28 Bar Outlet Temperature: 581.9 K Outlet Pressure: 35.93 Bar

Pressure Loss: 3.35 Bar

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Fuel Turbine Results

Substance: propanol Mass Flow Rate: 1.04 kg/s Inlet Temperature: 581.9 K Inlet Pressure: 35.93 Bar Outlet Temperature: 570.05 K Outlet Pressure: 26.51 Bar Temperature Loss: 11.86 K Pressure Loss: 9.42 Bar

Power: 10.21 kW

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Fuel Pipe 4 Results

Substance: propanol Mass Flow Rate: 1.04 kg/s Inlet Temperature: 570.05 K Inlet Pressure: 26.51 Bar Outlet Temperature: 570.05 K Outlet Pressure: 21.26 Bar

Pressure Loss: 5.25 Bar

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Oxidiser Line Component Outputs

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Oxidiser Pipe 1 Results

Substance: nitrous oxide Mass Flow Rate: 3.65 kg/s Inlet Temperature: 253.0 K Inlet Pressure: 25.0 Bar Outlet Temperature: 253.0 K Outlet Pressure: 24.87 Bar

Pressure Loss: 0.13 Bar

```

1. Any custom scripts using the package should be placed in and run from the main folder. testScript.py should serve as a useful guide for constructing your scripts.

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Example of Usage

SLRE has been designed to be (hopefully) quick and easy to use.

To begin, a thrust chamber object is created:

Python engineThrustChamber = ThrustChamber(fuelName='methane', oxName='oxygen', thrust=10*10**3, chamberPressure=40, fac=True, contractionRatio=3, ambientPressure=1.01325)

The combustion chamber and nozzle geometry for the thrust chamber is then generated, then combined into the overall thrust chamber geometry:

Python engineThrustChamber.getChamberGeometry(lStar=2.25, contractionLength=0.065, numberOfPointsConverging=100, numberOfPointsStraight=50) testThrustChamber.getRaoBellNozzleGeometry(0.8, numberOfPoints=100) testThrustChamber.getThrustChamberCoords()

This engine involves regenerative cooling so parameters for the solver are now defined:

Python solverParameters = SolverParameters(bartzEquationCoefficient=0.026, coolantSideHeatTransferCorrelation="sieder-tate")

To perform a regenerative cooling analysis, the cooling channels are defined, then fed into a regenerative cooling object along with the thrust chamber and solver paramters:

Python engineCoolingChannels = CoolingChannels(numberOfChannels=96, wallThickness=0.9e-3, midRibThickness=1.5e-3, channelHeight=1.25e-3, wallConductivity=365, wallRoughnessHeight=1e-6) engineRegenerativeCooling = RegenerativeCooling(thrustChamber=testThrustChamber, coolingChannels=testCoolingChannels, solverParameters=solverParameters)

To perform a cycle analysis, the state of the propellants in the tank are defined and Tank objects created to represent them:

```Python fuel = Propellant(testThrustChamber.fuel.name) ox = Propellant(testThrustChamber.ox.name)

fuel.defineState("T", 108, "P", 3105) ox.defineState("T", 60, "P", 310**5)

fuelTank = Tank(fuel) oxTank = Tank(ox) ```

Other components can be defined e.g:

```Python fuelPump = Pump(isentropicEfficiency=0.7, outletPressure=70e5) turbine = Turbine(isentropicEfficiency=0.7, outletPressure=testThrustChamber.injectionPressure10*5)

oxPump = Pump(isentropicEfficiency=0.7, outletPressure=testThrustChamber.injectionPressure10*5) ```

These can then be combined together to form their respective lines in the cycle:

Python fuelLine = Line(inletState=fuelTank.outletState, massFlowRate=engineThrustChamber.fuelMassFlowRate, components=[fuelPump, engineRegenerativeCooling, turbine]) oxLine = Line(inletState=oxTank.outletState, massFlowRate=thrustChamber.oxMassFlowRate, components=[oxPump])

The cycle is then formed from these lines:

Python cycle = Cycle(fuelLine=fuelLine, oxLine=oxLine, thrustChamber=engineThrustChamber)

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