Robert D. Bittner

Investigation of Scramjet Injection Strategies for High Mach Number Flows

A method for estimating the axial distribution of thrust performance potential in a supersonic combustor is described. A complementary technique for illustrating the spatial evolution and distribution of thrust potential and loss mechanisms in reacting flows is developed. A wall jet case and swept ramp injector case for Mach 17 and Mach 13.5 flight enthalpy inflow conditions, respectively, are numerically modeled and analyzed using these techniques.

Integrated Aeropropulsive Computational Fluid Dynamics Methodology for the Hyper-X Flight Experiment

Computational e uid dynamics tools have been used extensively in the analysis and development of the X-43A Hyper-X Research Vehicle. A signiecant element of this analysis is the prediction of integrated vehicle aeropropulsive performance,which includes an integration of aerodynamic and propulsion eoweelds. Thedevelopment of the Mach7X-43Arequiredapree ightassessmentoflongitudinalandlateral-directionalaeropropulsivecharacteristics nearthetargeteight-testcondition.Thedevelopmentofthispree ightdatabasewasaccomplishedthroughextensive aerodynamic wind-tunnel testing and a combination of three-dimensional inviscid airframe calculations and cowlto-tail scramjet cycle analyses togenerate longitudinal performance increments between mission sequences. These increments were measured directly and validated through tests of the Hyper-X e ight engine and vehicle e owpath simulator in the NASA Langley Research Center 8-Foot High Temperature Tunnel. Predictions were reened with tip-to-tailNavier‐Stokescalculations,which alsoprovidedinformation onscramjetexhaustplumeexpansion inthe aftbody region. A qualitative assessment of lateral-directional stability characteristics was made through a series of tip-to-tail inviscid calculations, including a simulation of the powered scramjet eight-test condition. Additional comparisons with wind-tunnel forceand momentdataas well as surfacepressuremeasurements from theHyper-X e ight engine and vehicle eowpath simulator model and wind-tunnel testing were made to assess solution accuracy.

Boundary Layer Transition on X-43A

The successful Mach 7 and 10 flights of the first fully integrated scramjet propulsion systems by the Hyper-X (X-43A) program have provided the means with which to verify the original design methodologies and assumptions. As part of Hyper-Xs propulsion-airframe integration, the forebody was designed to include a spanwise array of vortex generators to promote boundary layer transition ahead of the engine. Turbulence at the inlet is thought to provide the most reliable engine design and allows direct scaling of flight results to ground- based data. Pre-flight estimations of boundary layer transition, for both Mach 7 and 10 flight conditions, suggested that forebody boundary layer trips were required to ensure fully turbulent conditions upstream of the inlet. This paper presents the results of an analysis of the thermocouple measurements used to infer the dynamics of the transition process during the trajectories for both flights, on both the lower surface (to assess trip performance) and the upper surface (to assess natural transition). The approach used in the analysis of the thermocouple data is outlined, along with a discussion of the calculated local flow properties that correspond to the transition events as identified in the flight data. The present analysis has confirmed that the boundary layer trips performed as expected for both flights, providing turbulent flow ahead of the inlet during critical portions of the trajectory, while the upper surface was laminar as predicted by the pre-flight analysis.

Aerodynamic database development for the Hyper-X airframe integrated scramjet propulsion experiments

This paper provides an overview of the activities associated with the aerodynamic database which is being developed in support of NASAs Hyper-X scramjet flight experiments. Three flight tests are planned as part of the Hyper-X program. Each will utilize a small, nonrecoverable research vehicle with an airframe integrated scramj et propulsion engine. The research vehicles will be individually rocket boosted to the scramjet engine test points at Mach 7 and Mach l 0. The research vehicles will then separate from the first stage booster vehicle and the scramjet engine test will be conducted prior to the terminal decent phase of the flight. An overview is provided of the activities associated with the development of the Hyper-X aerodynamic database, including wind tunnel test activities and parallel CFD analysis efforts for all phases of the Hyper-X flight tests. A brief summary of the Hyper-X research vehicle aerodynamic characteristics is provided, including the direct and indirect effects of the airframe integrated scramjet propulsion system operation on the basic airframe stability and control characteristics. Brief comments on the planned post flight data analysis efforts are also included.

CFD Evaluation of Mach 17 HYPULSE Scramjet Combustor Data

Three-dimensional finite rate chemistry solutions are performed on a single fuel injector configuration. The results are compared with limited experimental data obtained from the HYPULSE expansion tube facility at simulated flight Mach 17 flow conditions. All comparisons, except for wall heat flux, were in excellent agreement. Key findings from this study are useful in interpretation of the experimental measurements.

Hypersonic CFD Applications for the National Aero-Space Plane

Design and analysis of the NASP depends heavily upon developing the critical technology areas that cover the entire engineering design of the vehicle. These areas include materials, structures, propulsion systems, propellants, integration of airframe and propulsion systems, controls, subsystems, and aerodynamics areas. Currently, verification of many of the classical engineering tools relies heavily on computational fluid dynamics. Advances are being made in the development of CFD codes to accomplish nose-to-tail analyses for hypersonic aircraft. Additional details involving the partial development, analysis, verification, and application of the CFL3D code and the SPARK combustor code are discussed. A nonequilibrium version of CFL3D that is presently being developed and tested is also described. Examples are given of portion calculations for research hypersonic aircraft geometries and comparisons with experiment data show good agreement.

Numerical study of external burning flowfields

This paper demonstrates the successful application of CFD to modeling an external burning flowfield. The study used the 2D, 3D, and PNS versions of the SPARK code. Various grids, boundary conditions, and ignition methodologies have been employed. Flameholding was achieved through the use of a subsonic outflow condition and a hot block located behind the step to ignite the fuel. Since the resulting burning produces a large subsonic region downstream of the cowl, this entire surface can be pressurized to the level of the back pressure. An evaluation of interactions between the ramjet exhaust and the external burning products demonstrate the complexity of this design issue. Ths code is now capable of evaluating the external burning effectiveness for flight vehicles using simple injector schemes, and the methodology can be readily applied to other external burning designs.