1. AAE450 Spring 2009 Lunar Lander Main Engine Thaddaeus Halsmer Thursday, February 12, 2009 Thaddaeus Halsmer, Propulsion

2. AAE450 Spring 2009 1.Main Engine Requirements (soft or semi-soft landing) Minimum Thrust = ~lunar lander weight at touchdown F min = 200 N Max Thrust restricted by engine throttling range F max = 2000 N High Isp Minimize Propellant Mass Dimensions Short to fit launch vehicle payload bay 2.Hydrogen Peroxide (H 2 O 2 ) / 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 H 2 O 2 – 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, Propulsion 1

3. AAE450 Spring 2009 Thaddaeus Halsmer, Propulsion 2 Polyethylene fuel plates Radial H 2 O 2 injection between fuel plates Nozzle Chamber Fig. 1 – Engine layout Fig. 2 – Dimensions D ch D noz L noz L ch Throttle Valve H2O2 Tank Feed lines M noz M ch + M fuel SymbolDescriptionValueUnits D ch chamber diameter0.5[m] D noz nozzle diameter0.13[m] L ch chamber length0.3[m] L noz nozzle length0.2[m] M ch mass of chamber minus ablative insulation, (graphite)4[kg] M fuel mass of fuel contained in chamber16.3[kg] M noz mass of nozzle13.5[kg] M eng engine mass = M noz + M ch + M fuel ~38[kg] M lines mass of feed lines & injectors???[kg] M tv mass of throttle valve???[kg] Table 1 – dimensions and masses

4. AAE450 Spring 2009 Thaddaeus Halsmer, Propulsion Backup slide (I)  EOM function for trajectory code, used to find prop mass 3 Fig. A1 – Trajectory EOM function

5. AAE450 Spring 2009 Thaddaeus Halsmer, Propulsion 4 Backup slide (II)  Trajectory code, used to find prop mass Fig. A2 – Trajectory function

6. AAE450 Spring 2009 Thaddaeus Halsmer, Propulsion Backup slide (III)  Engine performance analysis 5 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 flow rates of fuel and oxidizer can easily be found Fig. A3 – Sample CEA output

7. AAE450 Spring 2009 Thaddaeus Halsmer, Propulsion Backup slide (IV)  sizing equations and analysis methods 6 1.Engine sizing started with sizing fuel grain plates Used empirical value for fuel regression rate 1, 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