Artem A. Dyakonov

Aerothermodynamic Design of the Mars Science Laboratory Heatshield

Aerothermodynamic design environments are presented for the Mars Science Laboratory entry capsule heatshield. The design conditions are based on Navier-Stokes oweld simulations on shallow (maximum total heat load) and steep (maximum heat ux, shear stress, and pressure) entry trajectories from a 2009 launch. Boundary layer transition is expected prior to peak heat ux, a rst for Mars entry, and the heatshield environments were dened for a fully-turbulent heat pulse. The eects of distributed surface roughness on turbulent heat ux and shear stress peaks are included using empirical correlations. Additional biases and uncertainties are based on computational model comparisons with experimental data and sensitivity studies. The peak design conditions are 197 W=cm 2 for heat ux, 471 Pa for shear stress, 0.371 Earth atm for pressure, and 5477 J=cm 2 for total heat load. Time-varying conditions at xed heatshield locations were generated for thermal protection system analysis and ight instrumentation development. Finally, the aerothermodynamic eects of delaying launch until 2011 are previewed.

Aerothermodynamic Environments Definition for the Mars Science Laboratory Entry Capsule

An overview of the aerothermodynamic environments definition status is presented for the Mars Science Laboratory entry vehicle. The environments are based on Navier-Stokes flowfield simulations on a candidate aeroshell geometry and worst-case entry heating trajectories. Uncertainties for the flowfield predictions are based primarily on available ground data since Mars flight data are scarce. The forebody aerothermodynamics analysis focuses on boundary layer transition and turbulent heating augmentation. Turbulent transition is expected prior to peak heating, a first for Mars entry, resulting in augmented heat flux and shear stress at the same heatshield location. Afterbody computations are also shown with and without interference effects of reaction control system thruster plumes. Including uncertainties, analysis predicts that the heatshield may experience peaks of 225 W/sq cm for turbulent heat flux, 0.32 atm for stagnation pressure, and 400 Pa for turbulent shear stress. The afterbody heat flux without thruster plume interference is predicted to be 7 W/sq cm on the backshell and 10 W/sq cm on the parachute cover. If the reaction control jets are fired near peak dynamic pressure, the heat flux at localized areas could reach as high as 76 W/sq cm on the backshell and 38 W/sq cm on the parachute cover, including uncertainties. The final flight environments used for hardware design will be updated for any changes in the aeroshell configuration, heating design trajectories, or uncertainties.

Aerodynamic Challenges for the Mars Science Laboratory Entry, Descent and Landing

An overview of several important aerodynamics challenges new to the Mars Science Laboratory (MSL) entry vehicle are presented. The MSL entry capsule is a 70 degree sphere cone-based on the original Mars Viking entry capsule. Due to payload and landing accuracy requirements, MSL will be flying at the highest lift-to-drag ratio of any capsule sent to Mars (L/D = 0.24). The capsule will also be flying a guided entry, performing bank maneuvers, a first for Mars entry. The systems mechanical design and increased performance requirements require an expansion of the MSL flight envelope beyond those of historical missions. In certain areas, the experience gained by Viking and other recent Mars missions can no longer be claimed as heritage information. New analysis and testing is re1quired to ensure the safe flight of the MSL entry vehicle. The challenge topics include: hypersonic gas chemistry and laminar-versus-turbulent flow effects on trim angle, a general risk assessment of flying at greater angles-of-attack than Viking, quantifying the aerodynamic interactions induced by a new reaction control system and a risk assessment of recontact of a series of masses jettisoned prior to parachute deploy. An overview of the analysis and tests being conducted to understand and reduce risk in each of these areas is presented. The need for proper modeling and implementation of uncertainties for use in trajectory simulation has resulted in a revision of prior models and additional analysis for the MSL entry vehicle. The six degree-of-freedom uncertainty model and new analysis to quantify roll torque dispersions are presented.

Transition Onset and Turbulent Heating Measurements for the Mars Science Laboratory Entry Vehicle

An investigation of transitional/turbulent heating on the Mars Science Laboratory entry vehicle has been conducted. Laminar, transitional, and turbulent heating data were obtained in a perfect-gas, Mach 6 air wind tunnel and in a high-enthalpy shock tunnel in CO2. Flow field solutions were computed using a Navier-Stokes solver at the test conditions and comparisons were made between measured and predicted heating levels. Close agreement was obtained for all laminar perfect-gas cases. For the high-enthalpy CO2 cases, close agreement with the data was achieved when a fully-catalytic wall boundary condition was employed, whereas the predictions exceeded the data by more than 25% if a noncatalytic boundary condition was used. Turbulent heating predictions fell below the perfectgas air data by 25% but exceeded the CO2 data by 60%. Transition onset locations were determined through comparisons with laminar heating predictions, and boundary-layer parameters from the flow field solutions were used to develop correlations for the transition onset location and the turbulent heating augmentation on the leeside of the vehicle.

Aerodynamic Interference Due to MSL Reaction Control System

An investigation of effectiveness of the reaction control system (RCS) of Mars Science Laboratory (MSL) entry capsule during atmospheric flight has been conducted. The reason for the investigation is that MSL is designed to fly a lifting actively guided entry with hypersonic bank maneuvers, therefore an understanding of RCS effectiveness is required. In the course of the study several jet configurations were evaluated using Langley Aerothermal Upwind Relaxation Algorithm (LAURA) code, Data Parallel Line Relaxation (DPLR) code, Fully Unstructured 3D (FUN3D) code and an Overset Grid Flowsolver (OVERFLOW) code. Computations indicated that some of the proposed configurations might induce aero-RCS interactions, sufficient to impede and even overwhelm the intended control torques. It was found that the maximum potential for aero-RCS interference exists around peak dynamic pressure along the trajectory. Present analysis largely relies on computational methods. Ground testing, flight data and computational analyses are required to fully understand the problem. At the time of this writing some experimental work spanning range of Mach number 2.5 through 4.5 has been completed and used to establish preliminary levels of confidence for computations. As a result of the present work a final RCS configuration has been designed such as to minimize aero-interference effects and it is a design baseline for MSL entry capsule.

Development of Supersonic Retropropulsion for Future Mars Entry, Descent, and Landing Systems

Recent studies have concluded that Viking-era entry system deceleration technologies are extremely difficult to scale for progressively larger payloads (tens of metric tons) required for human Mars...

Aerothermodynamic Design of the Mars Science Laboratory Backshell and Parachute Cone

Aerothermodynamic design environments are presented for the Mars Science Laboratory entry capsule backshell and parachute cone. The design conditions are based on Navier-Stokes flowfield simulations on shallow (maximum total heat load) and steep (maximum heat flux) design entry trajectories from a 2009 launch. Transient interference effects from reaction control system thruster plumes were included in the design environments when necessary. The limiting backshell design heating conditions of 6.3 W/sq cm for heat flux and 377 J/sq cm for total heat load are not influenced by thruster firings. Similarly, the thrusters do not affect the parachute cover lid design environments (13 W/sq cm and 499 J/sq cm). If thruster jet firings occur near peak dynamic pressure, they will augment the design environments at the interface between the backshell and parachute cone (7 W/sq cm and 174 J/sq cm). Localized heat fluxes are higher near the thruster fairing during jet firings, but these areas did not require additional thermal protection material. Finally, heating bump factors were developed for antenna radomes on the parachute cone

Hypersonic and Supersonic Static Aerodynamics of Mars Science Laboratory Entry Vehicle

This paper describes the analysis of continuum static aerodynamics of Mars Science Laboratory (MSL) entry vehicle (EV). The method is derived from earlier work for Mars Exploration Rover (MER) and Mars Path Finder (MPF) and the appropriate additions are made in the areas where physics are different from what the prior entry systems would encounter. These additions include the considerations for the high angle of attack of MSL EV, ablation of the heatshield during entry, turbulent boundary layer, and other aspects relevant to the flight performance of MSL. Details of the work, the supporting data and conclusions of the investigation are presented.

Mars Science Laboratory Entry Capsule Aerothermodynamics and Thermal Protection System

The mars science laboratory (MSL) spacecraft is being designed to carry a large rover (> 800 kg) to the surface of Mars using a blunt-body entry capsule as the primary decelerator. The spacecraft is being designed for launch in 2009 and arrival at mars in 2010. The combination of large mass and diameter with non-zero angle-of-attack for MSL will result in unprecedented convective heating environments caused by turbulence prior to peak heating. Navier-Stokes computations predict a large turbulent heating augmentation for which there are no supporting flight data1 and little ground data for validation. Consequently, an extensive experimental program has been established specifically for MSL to understand the level of turbulent augmentation expected in flight. The experimental data support the prediction of turbulent transition and have also uncovered phenomena that cannot be replicated with available computational methods. The result is that the flight aeroheating environments predictions must include larger uncertainties than are typically used2 for a mars entry capsule. Finally, the thermal protection system (TPS) being used for MSL has not been flown at the heat flux, pressure, and shear stress combinations expected in flight, so a test program has been established to obtain conditions relevant to flight. This paper summarizes the aerothermodynamic definition analysis and TPS development, focusing on the challenges that are unique to MSL.

Development of Supersonic Retro-Propulsion for Future Mars Entry, Descent, and Landing Systems

1 ____________________________________________ * Aerospace Engineer, Atmospheric Flight & Entry Systems Branch, MS 489, Karl.T.Edquist@nasa.gov, Senior Member. † Aerospace Engineer, Atmospheric Flight & Entry Systems Branch, MS 489, Member. ‡ Propulsion Systems Engineer, Propulsion Systems Branch, MS EP4. § Senior Engineer, Aeroscience & Flight Mechanics Division, MS EG5, Member. ¶ Systems Engineer, Entry, Descent, and Landing Systems and Advanced Technologies Group, MS 321-220. # Research Scientist, Aerothermodynamics Branch, MS 230-2, Member. ** Aerospace Engineer, Systems Analysis Branch, MS 258-1, Member. †† Graduate Research Assistant, Daniel Guggenheim School of Aerospace Engineering, Student Member. Development of Supersonic Retro-Propulsion for Future Mars Entry, Descent, and Landing Systems