Ryan P. Starkey

Aerodynamic stability of reentry heat shield shapes for a crew exploration vehicle

A parametric study of the static stability of blunt-body reentry heat shield geometries applicable to a crew exploration vehicle has been performed. Performance trends are identified by varying geometric parameters that define a range of cross sections and axial shapes. Cross sections considered include oblate and prolate ellipses, rounded-edge polygons, and rounded-edge concave polygons. Axial shapes consist of the spherical segment, spherically blunted cone, and power law. Aerodynamic performance results that are based on a Newtonian surface pressure distribution have been verified against wind tunnel and flight data for the Apollo Command Module. Results are within 10% for aerodynamic coefficients, and trim angles of attack are computed within 1.2-deg. Stability and aerodynamic characteristics are observed to be more sensitive to changes in axial shape than changes in cross section. When uniform density is assumed, increased stability and performance are demonstrated at negative angles of attack for geometries with extremely blunt axial shapes and noneccentric cross sections. An unstable, oblate spherical segment at 20-deg angle of attack can produce a 56.1% increase in lift-to-drag ratio compared to a noneccentric stable spherical segment Shifting the center of gravity forward by 23.5% of its length can longitudinally stabilize this shield. The elliptical cross section, followed by the rounded-edge hexagon, and then by the rounded-edge concave hexagon rendered the most stable shapes.

Analytical Off-Design Lift-to-Drag-Ratio Analysis for Hypersonic Waveriders

An analytical, power-law-derived, waverider wing theory model is developed for studying the lift-to-drag-ratio characteristics of a rocket-powered waverider with a two-dimensional (planar) shock structure. Some inherent benee ts of the modeling method are explored, such as the decoupling of the vehicle length and planform shape fromtheothercomponentsinthelift-to-drag-ratio equation.Otherfactorsaffectinglift-to-dragratioarealsoinvestigated, including off-design angle-of-attack performance and sensitivity, off-design Mach number performance, altitude effects, base-pressure considerations, length-scale effects, and planform bluntness effects (i.e., spatulatevs caret-style waverider cone gurations ). The vehicle aerodynamics are derived in a manner that results in a similarity solution in which all results are independentof thewidth ofthevehicle.This similarity allowsfor easy vehicle scaling once a desirable cone guration has been determined.

Simple analytical model for parametric studies of hypersonic waveriders

An analytical, power-law derived hypersonic waverider model is developed using a three-dimensional wedge-based flowfield with viscous effects. The model is designed for simple parametric tradeoffs, and understanding of more complex optimization results. This analytical model is validated against a viscous optimized conical waverider created using an inverse design technique. Off-design performance of the analytical model at varying Mach numbers and angles-of-attack is validated using an optimized shock defined waverider (which itself has been validated with an Euler CFD calculation). Both the variable wedge angle waverider and shock defined waverider calculations had the same magnitude of error in comparison to the CFD validation L/Ds (maximum of 2% difference). The variable wedge angle method calculated the same aerodynamic and geometric properties of higher order methods in orders of magnitude less time.

Hypersonic Vehicle Telemetry Blackout Analysis

During certain hypersonic flight regimes, shock heating of air creates a plasma sheath, resulting in telemetry attenuation or blackout. The severity of the signal attenuation is dependent on vehicle configuration and orientation, flight trajectory, and transmission frequency. With the promise of airbreathing hypersonic vehicles looming on the horizon, telemetry solutions must be found to address safety considerations for flight testing (that is, flight termination and/or catastrophe analysis). This attenuation phenomena are investigated with a focus on the nonequilibrium plasma sheath properties (electron concentration, plasma frequency, collision frequency, and temperature) for a range of flight conditions and vehicle design considerations. Trajectory and transmission frequency requirements for airbreathing hypersonic vehicle design are then addressed, with comparisons made to both shuttle orbiter and Radio Attenuation Measurements C-II reentry flights. The effects of an applied magnetic field normal to ...

Sensitivity of hydrocarbon combustion modeling for hypersonic missile design

Aspects relating to the aerodynamic and propulsive design and analysis of missile-class, waverider-based hypersonic vehicles are explored in this paper. A quasi-one-dimensional engine model, including the effects of fuel injection, mixing, chemical production rates, heat transfer, and viscous losses, is developed and utilized to assess the effects of finite rate, hydrocarbon chemistry on optimized missile configurations. Resultant optimized single-and double-engine missile designs are shown for changes in fuel mixing length, fuel mixing efficiency, fuel-injector location, and assumed fuel mass fraction. The effects of these different design conditions on the cruise range are explored, as well as perturbations around these design points for optimized vehicles. Missiles are optimized for steady-state trim conditions at the beginning of cruise flight using parallelized genetic algorithm optimization software developed for this study. All missile designs are assumed to reach cruising altitude and velocity through the use of an external rocket booster. The missile is geometrically constrained to fit within the 0.61 x 0.61 × 4.27 m box limits for a naval vertical launch tube and has a desired cruise range of 750 km at Mach 8. Results show that the optimized combustor designs were extremely sensitive to small design perturbations. Two engine configurations are shown to be more robust than single-engine models for engine design perturbations.

Aerothermodynamic Optimization of Reentry Heat Shield Shapes for a Crew Exploration Vehicle

Gradient -based optimization of the aerodynamic perfor mance, static stability, and stagnation -point heat transfer has been performed to obtain optimal heat shield geometries for blunt -body planetary entry vehicles. Cross -sections considered include oblate and prolate ellipses, rounded -edge polygons, and round ed -edge concave polygons. Axial profiles consist of the spherical -segment, spherically -blunted cone, and power law. Aerodynamic models are based on modified Newtonian impact theory with semi -empirical shock -standoff distance and heat transfer correlations. Methods have been verified against wind tunnel and flight data of the Apollo Command Module and the Fire II experiment; they are within 12% for aerodynamic coefficients and stagnation -point heat fluxes. The selected design point corresponds to the setting in which the Apollo 4 Command Module generated its maximum heat flux, at an altitude of 61 km and a velocity of 10.3 km/s. Results indicate that oblate parallelogram configurations provide optimal sets of aerothermodynamic characteristics.

Significance of Nonequilibrium Surface Interactions in Stardust Return Capsule Ablation Modeling

The gas-surface modeling of high-density materials exposed to high-pressure atmospheric reentry conditions was extended to include low-density materials interacting with low-pressure atmospheric conditions. The fully implicit ablation thermal response code and multicomponent ablation thermochemistry program were extended to include nonequilibrium surface conditions for the Stardust return capsule. The Stardust return capsule reentered Earths atmosphere experiencing low pressure and had a low-density heat shield material. The material response of the Stardust return capsule was previously only modeled with surface equilibrium. Validation against Stardust reentry flight data showed that the equilibrium assumption led to an overprediction of recession and that the inclusion of nonequilibrium reduced the overprediction of this parameter. Incorporating the Park finite rate model nonequilibrium surface conditions led to a reduction in the calculated value of recession at the stagnation point from 1.12 to 0.72 cm in a nominal simulation ofthe Stardust return capsule. The nonequilibrium recession was closer to the measured recession of 0.65 cm. Incorporating nonequilibrium conditions also decreased the calculated total heat load from 28 to 19 kJ/cm 2 . The improved material response method is applicable to a range of reentries, including future missions such as the Orion crew exploration vehicle.

Off-Design Performance Characterization of a Variable Geometry Scramjet

Aspects relating to the design and analysis of scramjet engines for peak on and off-design performance are explored in this paper. A tip-to-tail Mach 8 flowpath is developed and compared using both fixed and variable geometry inlet and combustor components to assess the off-design performance. A quasi-one-dimensional scramjet engine model, including the effects of fuel injection location, mixing length, fuel mixing efficiency, chemical production rates, heat transfer, and viscous losses is utilized to assess the effects of the geometry variations on the hydrogen finite-rate chemistry. Due to the coupled nature of hypersonic propulsion/airframe integration, realistic two-dimensional inlet and nozzle designs have been included in this analysis. The scramjet characterization is done for changes in Mach number, dynamic pressure, angle of attack, and fuel equivalence ratio. The cowl leading edge position is used to track the trailing shock to enforce a shock-on-lip constraint, while the effects of changing the fuel injector location and the second expansion angle in the combustor are assessed. The off-design performance benefits/penalties of the selected geometric variations is quantified.

Aerogravity Assist Maneuvers: Coupled Trajectory and Vehicle Shape Optimization

The aero-gravity assist maneuver is proposed as a tool to improve the efficiency of the gravity assist as, due to the interaction with the planetary atmosphere, the angular deviation of the velocity vector can be definitely increased. Even though the drag reduces the spacecraft velocity, the overall Δv gain could be remarkable whenever a high lift-to-drag vehicle flies. A previous study addressed the 3D dynamic modeling and optimization of the maneuver including heliocentric plane change, heating rate, and structural load analysis. A multidisciplinary study of aero-gravity assist is proposed, focusing on coupled trajectory and vehicle shape optimization. A planar aero-gravity assist of Mars has been selected as a test case with the aim of maximizing the vehicle heliocentric velocity and limiting the heating rate experienced during the atmospheric path. A multiobjective approach has been adopted, and a particle swarm optimization algorithm has been chosen to detect the set of Pareto optimal solutions. The study includes a further refinement of the trajectory for three significant shapes belonging to the Pareto curve. The associated optimal control problem has been solved by selecting a direct method approach. The dynamics has been transcribed into a set of nonlinear constraints and the arising non linear programming problem has been solved through a sequential quadratic programming solver.

Axisymmetric Inlet Design for Combined Cycle Engines

Performance considerations for a turbine-based combined-cycle engine inlet are presented using the inlet of the Lockheed SR-71 as a baseline. The methodology incorporated is to modify the SR-71 inlet to expand its operational envelope to higher speeds. A numerical model is developed using the axisymmetric method of characteristics to perform full inviscid flow analysis, including any internal shock reflections. Starting characteristics are quantified based on the Kantrowitz limit. The original SR-71 inlet is analyzed throughout the designed flight regime, beginning at Mach 1.7 and ending with the shock-on-lip condition at Mach 3.2. A series of modifications are then considered for their ability to extend the range of the inlet into the hypersonic flight regime. The results show that two modifications, the widened shoulder centerbody and the variable cone with reextension, have the ability to remain started into the Mach 6‐7 range and have similar compressive performance to each other across the entire Mach spectrum. Nomenclature A = area M = Mach number m = mass flow P = pressure r = radial location T = temperature x = axial location γ = ratio of specific heats χ = flowfield property