Ramadas K. Prabhu

Entry Configurations and Performance Comparisons for the Mars Smart Lander

The Mars Smart Lander (MSL, renamed and redefined as the Mars Science Laboratory) will provide scientists with access to previously unachievable landing sites by providing precision landing to less than 10 km of a target landing site with landing altitude capability to 2.5 km above the Mars Orbiter Laser Altimeter geoid. Precision landing is achieved by using the aerodynamic forces on the entry body to aeromaneuver through the Martian atmosphere during the entry phase of flight. The entry body is designed to provide aerodynamic lift. The direction of the aerodynamic lift vector, defined by the vehicle bank angle, is commanded by the onboard entry guidance, to converge downrange and crossrange errors by parachute deploy, while meeting the parachute deploy constraints. Several approaches and entry body configurations for providing aerodynamic lift can be considered, including axisymmetric capsule configurations with offset c.g.s using ballast or packaging, aerodynamically shaped capsule-type configurations, and alternate configurations such as mid-lift-to-drag-ratio vehicles. The design considerations, entry configurations, and entry performance of the Mars Smart Lander are described.

Pressure Gradient Effects on Hypersonic Cavity Flow Heating

The effect of a pressure gradient on the local heating disturbance of rectangular cavities tested at hypersonic freestream conditions has been globally assessed using the two-color phosphor thermography method. These experiments were conducted in the Langley 31-Inch Mach 10 Tunnel and were initiated in support of the Space Shuttle Return-To-Flight Program. Two blunted-nose test surface geometries were developed, including an expansion plate test surface with nearly constant negative pressure gradient and a flat plate surface with nearly zero pressure gradient. The test surface designs and flow characterizations were performed using two-dimensional laminar computational methods, while the experimental boundary layer state conditions were inferred using the measured heating distributions. Three-dimensional computational predictions of the entire model geometry were used as a check on the design process. Both open-flow and closed-flow cavities were tested on each test surface. The cavity design parameters and the test condition matrix were established using the computational predictions. Preliminary conclusions based on an analysis of only the cavity centerline data indicate that the presence of the pressure gradient did not alter the open cavity heating for laminar-entry/laminar-exit flows, but did raise the average floor heating for closed cavities. The results of these risk-reduction studies will be used to formulate a heating assessment of potential damage scenarios occurring during future Space Shuttle flights.

X-33 Aerodynamic and Aeroheating Computations for Wind Tunnel and Flight Conditions

This report provides an overview of hypersonic Computational Fluid Dynamics research conducted at the NASA Langley Research Center to support the Phase II development of the X-33 vehicle. The X-33, which is being developed by Lockheed-Martin in partnership with NASA, is an experimental Single-Stage-to-Orbit demonstrator that is intended to validate critical technologies for a full-scale Reusable Launch Vehicle. As part of the development of the X-33, CFD codes have been used to predict the aerodynamic and aeroheating characteristics of the vehicle. Laminar and turbulent predictions were generated for the X 33 vehicle using two finite- volume, Navier-Stokes solvers. Inviscid solutions were also generated with an Euler code. Computations were performed for Mach numbers of 4.0 to 10.0 at angles-of-attack from 10 deg to 48 deg with body flap deflections of 0, 10 and 20 deg. Comparisons between predictions and wind tunnel aerodynamic and aeroheating data are presented in this paper. Aeroheating and aerodynamic predictions for flight conditions are also presented.

Experimental Hypersonic Aerodynamic Characteristics of Mars Surveyor 2001 Precision Lander with Flap

Aerodynamic wind-tunnel screening tests were conducted on a 0.029-scale model of a proposed Mars Surveyor 2001 Precision Lander (70-deg half-angle spherically blunted cone with a conical afterbody). The primary experimental objective was to determine the effectiveness of a single flap to trim the vehicle at incidence during a lifting hypersonic planetary entry. The laminar force and moment data, presented in the form of coefficients, and shock patterns from schlieren photography were obtained in the facilities of the NASA Langley Aerothermodynamic Laboratory for postnormal shock Reynolds numbers (based on forebody diameter) ranging from 2.637 x 103 to 92.35 × 10 3 , angles of attack ranging from 0 up to 23 deg at 0- and 2-deg sideslip, and normal-shock density ratios of 5 and 12. Based upon the proposed entry trajectory of the 2001 Lander, tests in the heavy gas CF 4 simulate a Mach number of approximately 12 based upon a normal shock density ratio of 12 in flight at Mars. The results from this experimental study suggest that when the traditional means of providing aerodynamic trim for this class of planetary entry vehicle are not possible (e.g., offset c.g.), a single flap can provide similar aerodynamic performance. An assessment of blunt-body aerodynamic effects attributed to a real gas was obtained by synergistic testing in Mach 6 ideal air at a comparable Reynolds number. From an aerodynamic perspective, an appropriately sized flap was found to provide sufficient trim capability at the desired lift-to-drag ratio for precision landing. Inviscid hypersonic flow computations using an unstructured grid were made to provide an assessment of the viability of a flap to provide aerodynamic trim to the Lander. Subsequent Navier-Stokes computational predictions were found to be in very good agreement with experimental measurement.

Aerodynamic Database Development for Mars Smart Lander Vehicle Configurations

An aerodynamic database has been generated for the Mars Smart Lander Shelf-All configuration using computational fluid dynamics (CFD) simulations. Three different CFD codes, USM3D and FELISA, based on unstructured grid technology and LAURA, an established and validated structured CFD code, were used. As part of this database development, the results for the Mars continuum were validated with experimental data and comparisons made where applicable. The validation of USM3D and LAURA with the Unitary experimental data, the use of intermediate LAURA check analyses, as well as the validation of FELISA with the Mach 6 CF4 experimental data provided a higher confidence in the ability for CFD to provide aerodynamic data in order to determine the static trim characteristics for longitudinal stability. The analyses of the noncontinuum regime showed the existence of multiple trim angles of attack that can be unstable or stable trim points. This information is needed to design guidance controller throughout the trajectory. Nomenclature A Reference area (m2) BC Ballistic coefficient (m /CD*A) CA Axial force coefficient