AAE450 Spring 2009

Lunar LanderMain Engine

Thaddaeus HalsmerThursday, February 12, 2009

Thaddaeus Halsmer, Propulsion

AAE450 Spring 2009

1. Main Engine Requirements (soft or semi-soft landing)• Minimum Thrust = ~lunar lander weight at touchdown

• Fmin = 200 N

• Max Thrust restricted by engine throttling range• Fmax = 2000 N

• High Isp• Minimize Propellant Mass

• Dimensions• Short to fit launch vehicle payload bay

2. Hydrogen Peroxide (H2O2) / Polyethylene Radial Flow Hybrid Engine• Similar performance to Bi-Prop systems but much less complex• Easily throttled by single valve• Isp ~295 s minimizing propellant mass

• Single propellant tank for H2O2 – solid fuel grain in combustion chamber

• Affordable ~$250,000 and 1 year to develop* (GLXP mission)• Scalable – build to suit, however larger engine will increase development cost

* Heister, S. D., (Communication, January 2009), Professor of Propulsion, Purdue University School of Aeronautics and Astronautics, Armstrong Hall Rm. 3331, West Lafayette, IN

Thaddaeus Halsmer, Propulsion1

AAE450 Spring 2009 Thaddaeus Halsmer, Propulsion2

Polyethylene fuel plates Radial H2O2 injection

between fuel plates

Nozzle

Chamber

Fig. 1 – Engine layout

Fig. 2 – Dimensions

Dch

Dnoz

Lnoz

Lch

Throttle Valve

H2O2 Tank

Feed lines

Mnoz

Mch + Mfuel

Symbol Description Value Units

Dchchamber diameter 0.5 [m]

Dnoznozzle diameter 0.13 [m]

Lchchamber length 0.3 [m]

Lnoznozzle length 0.2 [m]

Mchmass of chamber minus ablative insulation, (graphite) 4 [kg]

Mfuelmass of fuel contained in chamber 16.3 [kg]

Mnozmass of nozzle 13.5 [kg]

Mengengine mass = Mnoz + Mch+ Mfuel ~38 [kg]

Mlinesmass of feed lines & injectors ??? [kg]

Mtvmass of throttle valve ??? [kg]

Table 1 – dimensions and masses

AAE450 Spring 2009 Thaddaeus Halsmer, Propulsion

Backup slide (I) EOM function for trajectory code, used to find prop mass

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Fig. A1 – Trajectory EOM function

AAE450 Spring 2009 Thaddaeus Halsmer, Propulsion4

Backup slide (II) Trajectory code, used to find prop mass

Fig. A2 – Trajectory function

AAE450 Spring 2009 Thaddaeus Halsmer, Propulsion

Backup slide (III) Engine performance analysis

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1. Used CEA to analyze various chamber pressures and nozzle area ratios and compared output to historical values for sanity check

• Values used from CEA output• Isp = 295 s• C* = 1694.5• O/F = 8.08

• Using this equation: a known

thrust level and the O/F ratio, the mass flowrates of fuel and oxidizer can easily be found

Fig. A3 – Sample CEA output

gm

FIsp

AAE450 Spring 2009 Thaddaeus Halsmer, Propulsion

Backup slide (IV) sizing equations and analysis methods

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1. Engine sizing started with sizing fuel grain plates• Used empirical value for fuel regression rate1, rdot = 0.000305 m/s

• 1 Caravella, J.R., Heister, S. D., and Wernimont, E.J., “Characterization of fuel regression in a Radial Flow Hybrid Rocket,” Journal of Propulsion and Power, Vol. 14, No. 1, 1998, pp. 51-56.

• Using the fuel regression rate it and fuel density it was possible to define the surface area that provides the required fuel mass flow rate.

• Thickness of the fuel grain plates is determined from burn time and regression rate

2. Chamber mass was based on a chamber pressure of 17.2 Bar and Eq. 6.11 from Space Propulsion Analysis and Design (SPAD)

3. Nozzle size and mass was estimated using Eq. 3.133, 5.41 & 7.013 from SPAD