Search

The Mars 2020 entry vehicle's thermal protection system included measurements of in-depth temperature on the heatshield and backshell, and surface heating on the backshell. This paper describes the flight data and resulting analysis: subsurface temperatures at 11 heatshield and 6 backshell locations, measured total heat flux at two backshell locations, and a radiometer on the backshell to measure shock layer radiation. All temperatures were within allowable material limits. Total surface heat flux was derived from the in-depth temperature measurements at the thermocouple locations. Turbulent boundary-layer conditions occurred at nine heatshield locations and the radiometer provided evidence that heating from shock layer radiation significantly exceeded convective heating on the backshell. Computational results on the reconstructed entry trajectory are compared to the flight data. After boundary-layer transition, total heat flux from algebraic turbulence model calculations qualitatively matches the reconstructed heating, but with lower peak magnitudes. On the backshell, laminar heat flux predictions generally exceed the as-flown values by less than 1 W/cm², with shock layer radiation predicted to contribute the majority of heating. At the radiometer location, the maximum predicted radiative heat flux exceeds the measurement, with the measured value likely suppressed by ablation products on the sapphire window. [ABSTRACT FROM AUTHOR]

The Mars Science Laboratory heat shield was designed to withstand a fully turbulent heat pulse using information from ground testing and computational analysis on a preflight design trajectory. Instrumentation on the flight heat shield measured in-depth temperatures to permit reconstruction of the surface heating. The data indicate that boundary-layer transition occurred at five of seven measurement locations before peak heating. Data oscillations at three pressure measurement locations may also indicate transition. This paper presents the heat shield temperature and pressure data, possible explanations for the timing of boundary-layer transition, and a comparison of reconstructed and computational heating on the actual trajectory. A smooth-wall boundary-layer Reynolds number that was used to predict transition is compared with observed transition at various heat shield locations. A single transition Reynolds number criterion does not uniformly explain the timing of boundary-layer transition observed during flight. A roughness-based Reynolds number suggests that transition due to discrete or distributed roughness elements occurred. However, the distributed roughness height from acreage heat shield material would have needed to be larger than expected. The instrumentation confirmed the predicted location of maximum turbulent heat flux near the lee-side shoulder. The reconstructed heat flux at that location is bounded by smooth-wall turbulent convective heating, indicating that a significant augmentation due to surface roughness did not occur. Turbulent heating on the downstream side of the heat shield nose exceeded smooth-wall convective levels, assuming a supercatalytic surface, indicating that roughness may have augmented heating. The stagnation region heating also exceeded calculated convective heating; the cause of elevated heating may be attributed to a combination of shock-layer radiation and a heating augmentation of unknown origin that was also evident in ground test data. [ABSTRACT FROM AUTHOR]

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 exploration. Supersonic retropropulsion is one of a few developing technologies that may enable future human-scale Mars entry systems. However, in order to be considered as a viable technology for future missions, supersonic retropropulsion will require significant maturation beyond its current state. This paper proposes major milestones for advancing the component technologies of supersonic retropropulsion such that it can be reliably used on Mars technology demonstration missions to land larger payloads than are currently possible using Viking-based systems. The development roadmap includes technology gates that are achieved through ground-based testing and high-fidelity analysis, culminating with subscale flight testing in Earth's atmosphere that demonstrates stable and controlled flight. The component technologies requiring advancement include large engines (100s of kilonewtons of thrust) capable of throttling and gimbaling, entry vehicle aerodynamics and aerothermodynamics modeling, entry vehicle stability and control methods, reference vehicle systems engineering and analyses, and high-fidelity models for entry trajectory simulations. Finally, a notional schedule is proposed for advancing the technology from suborbital free-flight tests at Earth through larger and more complex system-level technology demonstrations and precursor missions at Mars. [ABSTRACT FROM AUTHOR]

The preflight aerodynamic database for the Mars Phoenix entry capsule is presented. The aerodynamic coefficients were generated as a function of total angle of attack and either Knudsen number, velocity, or Mach number, depending on the flight regime. The database was constructed using continuum flowfield computations and data from the Mars Exploration Rover and Viking programs. Hypersonic and supersonic static coefficients were derived from Navier-Stokes solutions on a preflight design trajectory. High-altitude data (free-molecular and transitional regimes) and dynamic pitch damping characteristics were taken from Mars Exploration Rover analysis and testing. Transonic static coefficients from Viking wind-tunnel tests were used for capsule aerodynamics under the parachute. Static instabilities were predicted at two points along the reference trajectory and were verified by reconstructed flight data. During the hypersonic instability, the capsule was predicted to trim at angles as high as 2.5 deg with an onaxis center of gravity. Peak trim angles for an offnominal pitching moment and a 5 mm offaxis center of gravity were predicted to be 4.2 and 4.8 deg, respectively. Finally, hypersonic static coefficient sensitivities to atmospheric density were predicted to be within uncertainty bounds. [ABSTRACT FROM AUTHOR]

The objective of this paper is to present the results of a detailed uncertainty analysis for high-fidelity fluid-structure interaction modeling of a deformable hypersonic inflatable aerodynamic decelerator at peak heating conditions for lifting Mars entry with a turbulent flow assumption. Uncertainty results are presented for the structural deformation response and surface conditions (pressure, shear stress, and convective heat transfer) of the inflatable decelerator with an efficient polynomial chaos expansion approach. The uncertainty results are compared with results obtained in a previous study for ballistic Mars entry. Approximately half of the flowfield and structural modeling uncertainties show at least 90% combined contribution to the inflatable decelerator deflection and resulting surface condition uncertainties. For lifting Mars entry, global nonlinear sensitivity analysis shows that the tensile stiffness of the inflatable structure's axial cords and radial straps and the torus torsional and tensile stiffnesses are the main contributors to the inflatable decelerator deflection uncertainty. As a result of these structural uncertainty contributions, the shape deformation contributes up to 10% of the uncertainty in the surface conditions. However, the freestream density dominates the uncertainty in the surface conditions experienced by the inflatable decelerator. In addition, the CO2-CO2 binary collision interaction is a significant contributor to aerodynamic heating and shear stress uncertainty. [ABSTRACT FROM AUTHOR]

This paper examines transient simulations performed to predict in-depth thermal response and surface recession of the proposed heat shield material for the Mars Science Laboratory entry capsule, that is, phenolic impregnated carbon ablator. The finite volume material response code used in this paper solves the time-dependent governing equations, including energy conservation and a three-component decomposition model, with a surface energybalance condition and a moving grid system to predict shape change due to surface recession. The predicted in-depth thermal response of heat shield material generally agrees well with the thermocouple data under various arejet conditions. Also, two-dimensional computations using aerothermal environment for Mars entry (derived from a proposed three-sigma trajectory) are performed around the heat shield shoulder region, where high heating occurs as the result of angle of attack. Parametric studies are conducted to examine the effects of carbon-fiber orientation, material properties, and surface recession on heat shield bondline temperature history. It is proved that the fiber orientation configuration of the baseline heat shield has the lowest maximum bondline temperature. [ABSTRACT FROM AUTHOR]

The methodology employed for the thermal design and margins assessment of the Mars Science Laboratory aeroshell thermal protection system is reviewed. A new thermal margins policy was developed in the course of this work that provides additional rigor over previous methods. Because of a late change of thermal protection materials from the heritage super lightweight ablator 561V to phenolic impregnated carbon ablator, the design of the heat shield followed a nontraditional path in which the flight thickness was selected based on a mass (rather than thermal) limit. The material switch was followed by detailed thermal analyses that demonstrated that the baselined thickness was sufficient to provide adequate thermal protection to the vehicle without violating design requirements during a 3- sigma worst-case entry condition. The backshell material thickness was also finalized before the thermal sizing was completed, and the resulting analysis showed that there was more than sufficient material on the backshell. The parachute cone cover and backshell interface plate were the only major thermal protection system elements that followed a standard design process. Thermal sizing was performed for acreage and special features on the cone cover and interface plate, and the hardware was manufactured according to those analyses. [ABSTRACT FROM AUTHOR]

Supersonic retropropulsion was experimentally examined in the Ames Research Center 9x7 Foot Supersonic Wind Tunnel at Mach 1.8 and 2.4. The model, previously designed for and tested in the Langley Research Center Unitary Plan Wind Tunnel at Mach 2.4,3.5, and 4.6, was a 5-in.-diam 70 deg sphere-cone forebody with a 9.55-in.-long cylindrical aftbody. The forebody was designed to accommodate up to four 4:1 area ratio nozzles, one on the model centerline and the other three on the half-radins spaced 120 deg apart. Surface pressure and flow visualization were the primary measurements. Including high-speed data to investigate the dynamics of the interactions between the bow and nozzle shocks. Three blowing configurations were tested with thrust coefficients np to 10 and angles of attack up to 20 deg. Results and observations from the test are provided. [ABSTRACT FROM AUTHOR]

The Interior Exploration using Seismic Investigations, Geodesy and Heat Transport (InSight) mission touched down in Elysium Planitia on 26 November 2018, becoming NASA's eighth successful entry, descent, and landing (EDL) at Mars. InSight inherited the successful 2008 Phoenix (PHX) EDL system, flying a nonspinning, ballistic trajectory with a 70 deg sphere-cone aeroshell (2.65 m diameter), a disk-gap-band parachute, and pulsed terminal descent and landing engines. InSight and Phoenix exhibited similar behavior, primarily in terms of trim attitude and an uninitiated roll reversal correlated with dynamic pressure, although the behaviors observed for InSight were more severe. For InSight, the initial clockwise roll and nonzero trim angle of attack were significant contributors to a short timeline, larger-than-predicted deceleration loads, and cross-track and up-track errors in landing location. The InSight and Phoenix reconstructions together indicate behavior more characteristic of a single, continuous instability region, suggesting that errors in the trim behavior from hypersonic nonequilibrium aerodynamics predictions along the dynamic pressure pulse may have a more substantial impact on nonspinning, ballistic entry vehicle performance. [ABSTRACT FROM AUTHOR]

A detailed uncertainty analysis for high-fidelity flowfield simulations over a fixed aeroshell of hypersonic inflatable aerodynamic decelerator scale for Mars entry is presented for fully laminar and turbulent flows at peak stagnationpoint heating conditions. This study implements a sparse-collocation approach based on stochastic expansions for efficient and accurate uncertainty quantification under a large number of uncertainty sources in the computational model. The convective and radiative heating and shear stress uncertainties are computed over the hypersonic inflatable aerodynamic decelerator surface and are shown to vary due to a small fraction of 65 flowfield and radiation modeling parameters considered in the uncertainty analysis. The main contributors to the convective heating uncertainty near the stagnation point are the CO2,-CO2, CO2-0, and CO-O binary collision interactions, freestream density, and freestream velocity for both boundary-layer flows. In laminar flow, exothermic recombination reactions are more important at the shoulder. The main contributors to radiative heating at the nose and flank were the CO2 dissociation rate and CO heavy-particle excitation rates, whereas the freestream density showed importance toward the shoulder. The CO2-CO2 interaction and freestream velocity and density control the wall shear stress uncertainty. [ABSTRACT FROM AUTHOR]

Early in the development of the Mars Science Laboratory thermal protection system on the heatshield, project management planned to use Lockheed Martin's Super Light Ablator in honeycomb as the ablative material based on successful use on previous Mars entry heatshields and on stagnation arcjet tests at heating rates beyond the design levels. Because this heatshield would be the first to experience combined turbulent flow and high shear environments as it entered the Mars atmosphere, tests were performed in various arcjet facilities on flat-plate, wedge, and swept-cylinder specimen configurations in order to ascertain the effects of shear on the material. During the course of these tests, a set of conditions within the flight envelope was identified that resulted in catastrophic failure in the SLA-561V. Consequently, project management decided to replace the SLA-561 V with the phenolic-impregnated carbon ablator, the material that had flown successfully on the Stardust mission and was undergoing intense testing and characterization for the Crew Exploration Vehicle Thermal Protection System Advanced Development Program. With only two years remaining before the expected launch date, and less than 18 months before the heatshield delivery date, the Mars Science Laboratory team developed and built NASA's first tiled ablator flight heatshield that contributed to the outstanding success of the spacecraft's entry, descent, and landing on 5 August 2012. [ABSTRACT FROM AUTHOR]

A validation study of computational fluid dynamics for supersonic retropropulsion was conducted using three Navier-Stokes flow solvers. The study compared results from the computational-fluid-dynamics codes to each other and to wind-tunnel test data obtained in the NASA Ames Research Center 9 x 7 ft Unitary Plan Wind Tunnel. Comparisons include surface pressure coefficient as well as unsteady plume effects and cover a range of Mach numbers, levels of thrust, and angles of orientation for zero-, one-, three-, and four-nozzle configurations. Flow-structure behavior changed with thrust and angle of orientation for all nozzle configurations. In general, the solvers compared best with the test data for the steadier cases of the one-nozzle and high-thrust three-nozzle configurations. Deviation in surface pressure was noted for the more unsteady cases and near transitions in behavioral modes. Strengths and weaknesses of the solvers are identified, and possible error sources are discussed. [ABSTRACT FROM AUTHOR]

Validation of computational fluid dynamics for supersonic retropropulsion is shown through the comparison of three Navier-Stokes solvers and wind-tunnel test results. The test was designed specifically for computational fluid dynamics validation and was conducted in the NASA Langley Research Center supersonic 4x4 foot Unitary Plan Wind Tunnel. The test includes variations in the number of nozzles, Mach and Reynolds numbers, thrust coefficient, and angles of orientation. Code-to-code and code-to-test comparisons are encouraging, and possible error sources are discussed. [ABSTRACT FROM AUTHOR]

Advancement of supersonic retropropulsion as a technology will rely heavily on the ability of computational methods to accurately predict vehicle aerodynamics during atmospheric descent, where supersonic retropropulsion will be employed. A wind-tunnel test at the NASA Langley Unitary Plan Wind Tunnel was specifically designed to aid in the support of Navier-Stokes codes for supersonic retropropulsion applications. Three computational fluid dynamics codes [data parallel line relaxation, fully unstructured Navier-Stokes three-dimensional, and overset grid flow solver] were exercised for multiple nozzle configurations for a range of freestream Mach numbers and nozzle thrust coefficients. The computational fluid dynamics pretest analysis of this wind-tunnel test aided in the test model design process by identifying the potential for tunnel blockage or unstart, of liquefaction within the plume, and of separation occurring at the internal fingers of the nozzles. This analysis led to a reduced model diameter, heating of the plenum, and reducing the nozzle area ratio, and the requirement to radius the corners at the fingers, to counter these potentials, respectively. Comparisons to test data were used to determine the existing capability of the codes to accurately model this complex flow, identify modeling shortcomings, and gain insight into the computational requirements necessary for correctly computing these flows. All three codes predict similar surface pressure coefficients and flowfield structures, such as jet termination shock, interface, bow shocks, and recirculation regions. However, the codes differ on the level of unsteadiness predicted. [ABSTRACT FROM AUTHOR]