Hypersonic Vehicle Systems Integration Vehicle Propulsion Integration and Force Accounting

1. Hypersonic Vehicle Systems IntegrationVehicle Propulsion Integration and Force Accounting 12 September, 2007 Dr. Kevin G. Bowcutt Senior Technical Fellow Chief Scientist of Hypersonics Boeing Phantom Works

2. Propulsion System Design & Integration Challenging Due to Aero-Propulsion Requirements and Complex Flow Physics

3. Hypersonic Vehicle Propulsion Requirements and Considerations

4. Approach to Aero-Propulsion Force Accounting Critical for Performance Analysis Ease, Accuracy and Consistency

5. Momentum Theory Extremely Valuable for Propulsion Analysis and Accurate Force Accounting Newton’s 2nd Law of Motion applied to a continuum fluid Draw a control volume, S, around a body, or through an engine flowpath from free stream to engine exit

6. First-Order Vehicle Propulsion Requirements Distinct propulsion systems are likely required for feasible and efficient and operation in different flight regimes (but may require close integration, i.e., “combined cycles”, to achieve needed synergy between propulsion systems) Subsonic through supersonic Hypersonic Exo-atmospheric Engine specific impulse or TSFC (i.e., engine efficiency) and thrust-to-weight ratio are primary propulsion parameters that drive design Fuel selection is a key driver of vehicle performance, size and cost Engine sizing is critical for achieving sufficient vehicle performance T/Wvehicle= 0.33 for takeoff T/Wvehicle= 0.25 for transonic and any other thrust pinch-points T/D = 3 across entire hypersonic flight regime

7. Hypersonic Air-breathing Propulsion Requires High Dynamic Pressure Flight to Generate Adequate Thrust

8. Engine Specific Impulse and Thrust-to-Drag Ratio Determine Propellant Consumption From Newton’s Second Law [F = (mV)], an equation can be derived for propellant fraction required between any two trajectory points: This equation can be used todetermine thrust / drag (T / D)requirements for reasonablepropellant mass fractions For realistic ?p (0.3 - 0.4),(T/D) = 3-4 required

9. Numerous Propulsion Technology Options Exist for Hypersonic Vehicles Turboramjets, ramjets/scramjets, rockets, combined-cycle engines

10. Typical Propulsion System Speed Limits

11. Low Speed Engine Choice Driven by Isp, Thrust-to-Weight & System Volume Tradeoffs Several low speed engine concepts are viable candidates Turbojet Turbo-ramjet Rocket Air augmented rocket (e.g., RBCC) Liquid air cycle engine Low speed engine generally sized for transonic pinch-point Large scramjet inlet and nozzle create high transonic drag

12. “Low-Speed” Engine Example: SR-71 J-58 Engine Operation Versus Subsonic Mach Number

13. “Low-Speed” Engine Example: SR-71 J-58 Engine Operation Versus Supersonic Mach Number

14. Turbine-Based Combined-Cycle (TBCC) Engine Mode Transition Challenge TBCC engines comprised of turbines mounted over ramjets/scramjets are promising propulsion systems for hypersonic cruise aircraft and reusable launch vehicles Feasibility of mode transition between engines has not yet been fully established Aerodynamic, mechanical and thermal interactions must be understood and managed Thrust margin must remain adequate with little or no operability risk Airframe-engine system must be able to tolerate and control events that could cause mission failure or vehicle loss (e.g., inlet unstart, engine flameout, thermal transients, etc.)

15. Other TBCC Engine Challenges Thermal issues of Mach 3-4 turbine engines Lightweight, durable, high-temperature materials High-temperature bearings, bushings, seals and wear surfaces Thermal management during operation and after shutdown Transonic and high Mach thrust performance Thrust pinch points/sufficient thrust over entire speed range Transonic nozzle drag & download of large scramjet nozzle Ground testing: verification and validation Facility size and Mach number/enthalpy constraints to accommodate both flowpaths of a TBCC engine Integrating turbine inlets and nozzles in complex dual-mode scramjet flowpath shapes Strong aerodynamic interactions between highly integrated dual inlets and dual nozzles

16. Subsonic and Supersonic Turbine Engine Performance Can Be Estimated via Simulation Tools NASA Glenn’s EngineSim 1.7a can be used to estimate turbofan, turbojet and ramjet engine weight, thrust and efficiency, as a function of speed and altitude (http://www.grc.nasa.gov/WWW/K-12/airplane/ngnsim.html) Engine size, and component design, material and performance parameters can be varied to match existing engines

17. Supersonic Combustion Ramjets (scramjets) Provide for Efficient Hypersonic Propulsion

18. Vehicle Forebody Considerations Driven by Requirements of Different Speed Regimes Low Speed Flow spillage drag Inlet starting Flow separation Inlet stability (unstart) Compression efficiency Combustor isolation High Speed Inlet mass capture Compression efficiency Optimum inlet contraction ratio Boundary layer transition Lateral flow spillage Inlet flow distortion Flow separation Corner flow effects Aerodynamic heating & insulation/cooling requirements Shock interactions with leading edges Optimum inlet axial location Longitudinal vehicle stability Pitch / yaw sensitivity

19. Forebody Requirements and Design Options Requirements High compression efficiency Low drag and high mass capture (high wair / D) Flow uniformity to inlet Minimal crossflow Uniform boundary layer transition Good volumetric efficiency Maintain aerodynamic center aft for stability Design Options Cones Wedges Ogives Spatular bodies Bielliptic cross-sections (lenticular) Waveriders Inward turning / Busemann inlets

20. Hypersonic Engine Thrust-to-Weight Can Be a Function of Vehicle Forebody Type Assumptions: Thrust ~ capture area Engine length equal for all cases Engine weight ~ engine internal wetted area ~ inlet perimeter Engine T/W ~ capture area / inlet perimeter Mitigating Factors: Fuel-air mixing difficulty increases with H May require longer engine or more intrusive fuel injectors (i.e., weight and drag) Forebody wetted area differences incur structural weight impacts

21. Inlet Air Capture Requirement VariesDramatically With Vehicle Mach Number Shock-on-lip at design (maximum) Mach number 14° cone external compression (drawn to scale)

22. Capture Area Drives Thrust-to-Drag RatioCapture Area Therefore a First-Order Configuration Driver The rocket equation be used to estimate vehicle capture area requirement, hence relative engine size Solve for A? / Axs from equation for T/D

23. Inlet Design Tradeoffs Driven By Inlet Performance, Drag, Weight and Complexity Inlet design options include external, internal and mixed compression concepts Each has different performance, drag, weight and complexity characteristics

24. Inlet Type and Shock Number Trends With Maximum Mach Number

25. Supersonic (Turbine) Inlet Design Layout Approach Select capture area size Select desired number of ramps and ramp angles At design Mach number, position ramps so that all ramp shocks focus on the cowl leading edge Another option is not focusing the shocks at the design condition, but then there will always exist supersonic spillage drag for Mach numbers at least as high as the design Mach number Note: For fully variable ramps and cowl, spillage drag can be made zero for all Mach numbers, but will likely result in increased cowl drag at supersonic speeds

26. Turbine Inlet Pressure Recovery Estimation Turbine inlet pressure recovery can be estimated from MIL standard curve (MIL-E-5008B), or trends from existing analytical predictions or actual airplane inlet data

27. Inlet Performance and Operational Issues / Requirements at Hypersonic Speeds High compression efficiency Inviscid (turn angles and shock strengths) Friction and heat loss (pressure distribution, boundary layer transition and wetted area) Viscous phenomena may affect performance and operability Entropy swallowing effects on boundary layer transition Shock / boundary layer interactions (potential flow separation) Corner flows Note: High air total temperature prevents use of boundary layer bleed to compensate for viscous effects. Actively cooled surfaces Required when convective heating rates exceed threshold Leading edge heating and drag Edges must be blunt to reduce heating and accommodate active cooling Shock / leading edge interactions increase heating dramatically Optimum axial location Influences inlet capture and / or performance, vehicle trim and c.g.

28. Hypersonic/Scramjet Inlet Design Layout Approach Select a cowl lip axial station: 0.45 – 0.55 times body length is reasonable Select desired number of inlet ramps, and whether cowl and/or ramps are variable At design Mach number, position cowl vertically so that forebody shock is on the cowl leading edge, then position ramps so that all ramp shocks focus on the cowl lip Another option is to not focus the shocks at the design condition, but then there will always exist supersonic spillage drag for Mach numbers at least as high as the design Mach number

29. Total Flow Turning in Optimum 3-Shock, 4-Shock, and Isentropic Inlets

30. Accounting for All Sources of Inlet Drag Critical to Accurate Force Accounting Inlet Spillage Drag: Additive + Incremental Cowl Drag (or Thrust) due to spilling flow Bypass Drag (due to flow momentum loss) Bleed Drag (due to flow momentum loss) Secondary Airflow Drag (due to flow momentum loss)

31. Effect of Forebody Boundary Layer Transition on Inlet Performance

32. High-Speed Propulsion Requirements Impact Transonic Aero Performance Inlet drag Body ramps required for high speed inlet efficiency, but . . . Ramps produce high spillage drag at transonic and low supersonic speeds Nozzle drag Large base area required for high speed thrust, but . . . Low nozzle pressure ratios result in aftbody flow separation, drag and large pitching moments (trim drag)

33. Transonic Inlet Drag The forebody ramp and inlet system of air-breathing hypersonic vehicles typically generate large drag at transonic conditions Large ramps required for adequate and efficient high-speed compression Inlet throat area must be small at high speeds to meet contraction ratio requirements for good engine performance Impact: Maximum throat size at low speeds is limited by mechanical (variable geometry) constraints and fuel penetration across flowpath duct

34. Nozzle Considerations vs. Speed Regime Subsonic/Transonic Overexpansion and drag (internal and external) Drag reduction Freestream interactions Combustor interactions(e.g., thermal choke location and orientation) Flow separation Ground effects Hypersonic Optimum balance of losses Divergence Friction drag Underexpansion Finite-rate chemistry (kinetics) Freestream interactions (vertical and lateral) Heat transfer and cooling Combustor flow profile effects Viscous interactions Relaminarization Contribution to vehicle lift and trim

35. Nozzle Expansion Area Requirement Varies Dramatically With Mach Number

36. Transonic Nozzle Drag Can Be Quite Large Nozzle region of space plane designs typically generate large drag at transonic conditions Engines produce relatively low nozzle pressure ratios Large areas required for high Mach scramjet efficiency Confirmed by experiment

37. Over-expanded Nozzle Base Drag First-Order Analysis Approach First determine fraction of nozzle expansion area filled by engine exhaust flow (i.e., area after flow expands to ambient pressure) Apply base pressure to remaining nozzle area

38. External Burning – An Approach To Transonic Nozzle Drag Reduction Injecting and burning fuel external to the engine cowl can fill nozzle base region with hot, low density gases at ambient pressure to reduce or eliminate nozzle drag

39. Optimum nozzle performance requires a careful balance between loss mechanisms across entire Mach range Low divergence losses (e.g., perfect nozzle) result in excessive viscous losses due to long lengths Gradual expansion to reduce kinetic losses results in large viscous and divergence losses Typical values of losses at hypersonic speeds Scramjet Nozzle Efficiency and Loss Mechanisms

40. Nozzle Design Variables Can Be Used to Tailor Nozzle Thrust, Lift and Pitching Moment