Orbital Payload Delivery Using Hydrogen and Hydrocarbon Fuelled Scramjet Engines M. R. Tetlow and C.J. Doolan School on Mechanical Engineering The University of Adelaide

Overview Current launch systems Scramjet background Mission profile and vehicle description Software operation Trajectory outputs Analysis of results Conclusions

Aim Design a mission using a hydrocarbon powered (JetA) and a hydrogen powered scramjet stage to reach a 200km circular orbit Compare the mission profiles and performance of the two launch systems Compare the performance to current rocket powered systems

Current Launch Systems Launch Vehicle Payload mass (mass fraction) LEO orbit and inclination ASLV 150kg (0.36%) 400km at 43° M-3S11 780kg (1.26%) 185km at 31° Long March CZ1D 720kg (0.9%) 200km at 28° Start-1 360kg (0.6%) 400km at 90° 1% at 200km is indicative of the performance of this class of vehicle [Isakowitz - 1995]

Scramjets Supersonic combustion ramjet Hydrogen fuelled Geometry dependent on operating conditions Hydrogen fuelled High energy, low storage density Operating range: Mach 5 to 15 Isp ~ 3000s Hydrocarbon fuelled Lower energy, high storage density Operating range: Mach 5 to 10 Isp ~1200s Minimum dynamic pressure ~10kPa

Waveriders

Waveriders Blended wing vehicle with integrated propulsion system “Ride” the shock wave Aerodynamics are Mach No. dependent Fuel mass fractions ε=0.58 for hydrogen fuelled vehicle ε=0.7 for hydrocarbon fuelled vehicle ε=0.9 for rockets

Quasi-1D Scramjet Propulsion Model Flow From Inlet Displacement Thickness Growth Combustion Area Change Shear Stress Ignition Delay Heat Transfer Injector

Quasi-1D Scramjet Propulsion Model Set of ODEs used to describe scramjet propulsion. 2-step chemistry model. Skin friction and wall heat transfer included. H2 and Jet A fuel options. Idealised hypersonic inlet (with losses) used to supply combustor. Lawrence Livermore ODEPACK Solver used for ODE solution.

T4 Experiment Parallel Combustor Scramjet model validated against shock tunnel data (T4, University of Queensland, Boyce et al., 2000). Parallel and diverging combustor data used for validation study. Good agreement obtained using an 88% combustion efficiency. A conservative 50% combustion efficiency was used for trajectory modelling (for combustor losses). Diverging Combustor

Common Design Parameters GLOW 9300kg 2 stage solid rocket booster Stage 1: 2420kg start mass, 1980kg propellant Stage 2: 4880kg start mass, 4000kg propellant Cranked wing concept with aerodynamics taken from a NASA study Rocket powered upper stage with performance based on the H2 upper stage

Software Models Simulation environment Target/constraints 3DOF dynamics model, rotating spheroidal earth model, 4th order gravitation model, MSISE 93 atmosphere model Target/constraints Velocity stopping condition Altitude and flight path angle targets for scramjet burn only Parameterised vertical acceleration profile

Common Mission Parameters Booster 1 burn 10s burn, Alt =9.5km, Vel =550m/s Coast 45.4s, Alt =15.9km, Vel =295m/s Booster 2 burn 25s burn, Alt =19.6km, Vel =2411m/s 44.6s, Alt =25.3km, Vel =2000m/s -------------- Orbital stage Two burns, Alt =200km, Vel =7784m/s

Mission Profiles for Hydrogen Fuelled Vehicle

Hydrogen Case - Altitude Profile

Hydrogen Case - Velocity Profile

Mission Profiles for Hydrocarbon Fuelled Vehicle

Hydrocarbon Case - Altitude Profile

Hydrocarbon Case - Velocity Profile

Payload Estimation Mass and state at end of the scramjet burn Scramjet mass fractions Hydrogen fuelled waverider εpropellant = 0.58 Hydrocarbon fuelled waverider εpropellant = 0.7 Orbital stage upper stage εstructure = 0.15 ΔV requirement based on Hohmann transfer

Mass Breakdown Hydrogen fuelled case Initial mass: 2000kg Fuel mass: 316kg Structure mass: 1000kg Orbital stage mass: 684kg Payload to 200km circular: 108.5kg Payload mass fraction: 1.16% Hydrocarbon fuelled case Initial mass: 2000kg Fuel mass: 258.8kg Structure mass: 918.3kg Orbital stage mass: 822.9kg Payload to 200km circular: 36kg Payload mass fraction: 0.38%

Discussion Payload mass fractions similar to rockets even though much higher Isp? Considerably lower fuel mass fractions i.e. more of stage mass is structure, compared to rockets Structure is more expensive than fuel. These systems need to be reusable to be financially viable

Discussion Lighter scramjet stage for the hydrocarbon fuelled system Hydrogen fuelled vehicle considerably higher payload capability than hydrocarbon fuelled case Longer duration burn at higher Isp for H2 case Better packing efficiency does not help the hydrocarbon vehicle as a large aerodynamic area is needed to maintain lift at high altitude so the vehicle cannot be made smaller.

Conclusions Similar payload mass fractions to rockets Therefore need to be reusable Hydrocarbon fuelled case has lighter structure than hydrogen fuelled case Better packing efficiency Better packing efficiency cannot be utilised due to aerodynamic requirements

Questions?