Christopher E. Glass

X-34 Experimental Aeroheating at Mach 6 and 10

Critical technologies are being developed to support the goals of the NASA Office of Aeronautics and Space Transportation Technology Access to Space initiative for next-generation reusable space transportation systems. From the perspective of aerothermodynamic performance throughout the flight trajectory, the Reusable Launch Vehicle program incorporates conceptual analysis, ground-based testing, and computational fluid dynamics to provide flyable suborbital flight demonstrator vehicles. This report provides an overview of the hypersonic aeroheating wind tunnel test program conducted at the NASA Langley Research Center in support of one of these vehicles, the X-34 small reusable technology demonstrator program. Global surface heat transfer images, surface streamline patterns, and shock shapes were measured on 0.0153- and 0.0183-scale models of proposed X-34 flight vehicles at Mach 6 and 10 in air. The primary parametrics that were investigated include angles-of-attack from 0 to 35 deg. and freestream unit Reynolds numbers from 0.5 to 8 million per foot (which was sufficient to produce laminar, transitional, and turbulent heating data), both with and without control surface deflections. Comparisons of the experimental data to computational predictions are included, along with a discussion of the implications of some of the experimental flow features for the flight vehicle.

Fluorescence Imaging of Underexpanded Jets and Comparison with CFD

*† ‡ § An experimental study of underexpanded and highly underexpanded axisymmetric nitrogen free jets seeded with 0.5% nitric oxide (NO) and issuing from a sonic orifice was conducted at NASA Langley Research Center. Reynolds numbers based on nozzle exit conditions ranged from 770 to 35,700, and nozzle exit-to-ambient jet pressure ratios ranged from 2 to 35. These flows were non-intrusively visualized with a spatial resolution of approximately 0.14 mm x 0.14 mm x 1 mm thick and a temporal resolution of 1µs using planar laser-induced fluorescence (PLIF) of NO, with the laser tuned to the stronglyfluorescing UV absorption bands of the Q1 band head near 226.256 nm. Three laminar cases were selected for comparison with computational fluid dynamics (CFD). The cases were run using GASP (General Aerodynamic Simulation Program) Version 4. Comparisons of the fundamental wavelength of the jet flow showed good agreement between CFD and experiment for all three test cases, while comparisons of Mach disk location and Mach disk diameter showed good agreement at lower jet pressure ratios, with a tendency to slightly underpredict these parameters with increasing jet pressure ratio.

ANALYSIS OF EFFECTIVENESS OF PHOENIX ENTRY REACTION CONTROL SYSTEM

Interaction between the external flowfield and the reaction control system (RCS) thruster plumes of the Phoenix capsule during entry has been investigated. The analysis covered rarefied, transitional, hypersonic and supersonic flight regimes. Performance of pitch, yaw and roll control authority channels was evaluated, with specific emphasis on the yaw channel due to its low nominal yaw control authority. Because Phoenix had already been constructed and its RCS could not be modified before flight, an assessment of RCS efficacy along the trajectory was needed to determine possible issues and to make necessary software changes. Effectiveness of the system at various regimes was evaluated using a hybrid DSMC-CFD technique, based on DSMC Analysis Code (DAC) code and General Aerodynamic Simulation Program (GASP), the LAURA (Langley Aerothermal Upwind Relaxation Algorithm) code, and the FUN3D (Fully Unstructured 3D) code. Results of the analysis at hypersonic and supersonic conditions suggest a significant aero-RCS interference which reduced the efficacy of the thrusters and could likely produce control reversal. Very little aero-RCS interference was predicted in rarefied and transitional regimes. A recommendation was made to the project to widen controller system deadbands to minimize (if not eliminate) the use of RCS thrusters through hypersonic and supersonic flight regimes, where their performance would be uncertain.

Comparison of a 3-D CFD-DSMC Solution Methodology With a Wind Tunnel Experiment

A solution method for problems that contain both continuum and rarefied flow regions is presented. The methodology is applied to flow about the Mars Sample Return Orbiter (MSRO), a 3‐D blunt body, which contains a region of highly compressed forebody flow, a shear layer where the flow separates from a forebody lip (i.e., where the base plane intersects the forebody), and a low density wake region about the afterbody. Because the blunt body flow contains such disparate regions, employing a single numerical technique to solve the entire flow field for a 3‐D configuration is often impractical, or the technique does not apply. Direct simulation Monte Carlo (DSMC) could be employed to solve the entire flow field; however, the technique would require inordinate computational resources for continuum to near continuum regions, since, it is best suited for the rarefied wake region. Computational fluid dynamics (CFD) is appropriate for the high‐density region on the forebody, but in the rarefied wake region, contin...

Airframe Research and Technology for Hypersonic Airbreathing Vehicles

The Hypersonics Investment Area (HIA) within NASAs Advanced Space Transportation Program (ASTP) has the responsibility to develop hypersonic airbreathing vehicles for access to space. The Airframe Research and Technology (AR and T) Project, as one of six projects in the HIA, will push the state-of-the-art in airframe and vehicle systems for low-cost, reliable, and safe space transportation. The individual technologies within the project are focused on advanced, breakthrough technologies in airframe and vehicle systems and cross-cutting activities that are the basis for improvements in these disciplines. Both low and medium technology readiness level (TRL) activities are being pursued. The key technical areas that will be addressed by the project include analysis and design tools, integrated vehicle health management (IVHM), composite (polymer, metal, and ceramic matrix) materials development, thermal/structural wall concepts, thermal protection systems, seals, leading edges, aerothermodynamics, and airframe/propulsion flowpath technology. Each of the technical areas or sub-projects within the Airframe R and T Project is described in this paper.

Numerical/Experimental Investigation of 3-D Swept Fin Shock Interactions

Three-dimensional forward swept fin shock-shock interactions are examined using laminar computational fluid dynamics. The results are compared to schlieren images and stagnation line heat flux data on a 0.5 inch diameter cylindrical edge fin previously published by Berry and Nowak. The cases investigated include Type III and Type IV shock interactions, with a high localized heat flux due to jet or shear layer impingement. Numerical results are presented for three shock-on-fin interactions with a fin leading edge sweep of zero (perpendicular to the freestream), 15°, and 25°. Results indicate that when the supersonic jet impinges on the body the resulting flow is unsteady due to vortex motion in the impingement region. The predicted peak in surface heat flux for the 15° forward swept case is very narrow, and is higher than the experimental data. The predicted peak for the 25° forward swept case is wider, and is lower than the experiment. Off-axis heat flux was not measured experimentally, but the computations show that circumferential heat flux gradients are an order of magnitude smaller than longitudinal. The peak heat flux falls by only 5% in the first 11.5° around the fin, indicating that off-axis heating is an important design consideration in these interactions.

PHARO—Propellant harvesting of atmospheric resources in orbit

Collection and storage of propellant on-orbit has the potential to dramatically reduce launch mass for future exploration missions.12 A proposed method for this collection utilizes an orbiting vehicle that collects ambient air at a high altitude and uses a fraction of the air for orbital maintenance while storing the remainder for exploration propellant. The derivation of the relations governing propulsion requirements of thrust and specific impulse is presented. Initial requirements for the collector are defined through design maps based on a notional Mars mission.

Analysis of effectiveness of Phoenix Entry Reaction Control System

Interaction between the external flowfield and the reaction control system (RCS) thruster plumes of the Phoenix capsule during entry has been investigated. The analysis covered rarefied, transitional, hypersonic and supersonic flight regimes. Performance of pitch, yaw and roll control authority channels was evaluated, with specific emphasis on the yaw channel due to its low nominal yaw control authority. Because Phoenix had already been constructed and its RCS could not be modified before flight, an assessment of RCS efficacy along the trajectory was needed to determine possible issues and to make necessary software changes. Effectiveness of the system at various regimes was evaluated using a hybrid DSMC-CFD technique, based on DSMC Analysis Code (DAC) code and General Aerodynamic Simulation Program (GASP), the LAURA (Langley Aerothermal Upwind Relaxation Algorithm) code, and the FUN3D (Fully Unstructured 3D) code. Results of the analysis at hypersonic and supersonic conditions suggest a significant aero-RCS interference which reduced the efficacy of the thrusters and could likely produce control reversal. Very little aero-RCS interference was predicted in rarefied and transitional regimes. A recommendation was made to the project to widen controller system deadbands to minimize (if not eliminate) the use of RCS thrusters through hypersonic and supersonic flight regimes, where their performance would be uncertain.

Hypersonic Shock Interactions About a 25°/65° Sharp Double Cone

This paper presents the results of a numerical study of shock interactions resulting from Mach 10 air flow about a sharp double cone. Computations are made with the direct simulation Monte Carlo (DSMC) method by using two different codes: the G2 code of Bird and the DAC (DSMC Analysis Code) code of LeBeau. The flow conditions are the pretest nominal free‐stream conditions specified for the ONERA R5Ch low‐density wind tunnel. The focus is on the sensitivity of the interactions to grid resolution while providing information concerning the flow structure and surface results for the extent of separation, heating, pressure, and skin friction.