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A parametric structural investigation of a new atmospheric entry vehicle concept for future planetary probes and sample return vehicles was conducted. During launch and reentry, the vehicle may experience large structural loads. To enable rapid assessment of candidate vehicle configurations for different missions, a parametric structural dynamics analysis methodology was developed. The methodology generates finite-element models and enables rapid modification to all aspects of the models, including geometry, material properties, load and boundary conditions, mesh parameters, and analysis controls. Analyses including quasi-static inertial loading, vibration frequency response, acoustic loading, and reentry loading were conducted for the structural concept study. Several geometry, mass, and material parameters were varied from -5 0 to +50% of their baseline value and analyzed. A total of 136 analyses were conducted on 34 unique vehicle configurations. Of the parameters investigated, cone angle and vehicle diameter were found to be the most influential on the vehicle's mass and structural response. The highest stress observed was for an inertial launch analysis conducted on a -50% cone angle model, which had 2.7x the baseline stress and only 1.4x the baseline mass. This analysis methodology is shown to facilitate and expedite future planetary exploration mission analyses. [ABSTRACT FROM AUTHOR]

Future planetary exploration requires spacecraft to land softly on rough terrain and in severe environments. Since conventional landing methods have problems such as high rebounds and excessive resource consumption, the baseextension separation mechanism, which combines springs and separable units, is proposed as a novel landing mechanism. Although the mechanism performed good soft landings, the performance evaluation was limited. Therefore, this study evaluated its performance multilaterally. The proposed technology was analytically compared with two other landing technologies: a generalized-hybrid momentum exchange impact damper and an aluminum foam landing gear. The proposed technology suppressed rebound and acceleration better than the generalized momentum exchange impact damper. Once the components of the proposed technology had been lightened, its energy conversion efficiency matched that of the aluminum foam landing gear. In addition, experiments were conducted using small-scale models to confirm the feasibility. The experiments showed that the proposed technology has good soft landing performance, matching the results of the simulations. Finally, design optimization was discussed for further performance improvements. It was clarified that both the spring stroke and the pre-tension length of the spring should be large. Optimizing the design of the spring improved its performance and compensated for its disadvantages. [ABSTRACT FROM AUTHOR]

In this paper, the rolling stability of a power-generating tumbleweed rover is investigated. The nonlinear equations of motion of the rover capture the nonholonomie nature of rolling without slipping, the external wind force, and energy dissipation due to rolling friction and power generation. To assess stability of the system as it rolls about a preferred axis of rotation, Lyapunov's indirect method is employed. When the shell of the tumbleweed rover has a uniform mass distribution, Lyapunov's indirect method predicts instability; when the inertia about a preferred axis of rotation is increased, Lyapunov's indirect method predicts stability. Numerical simulation results of both the linearized and nonlinear equations of motion confirm both the instability and stability predictions of Lyapunov's indirect method. [ABSTRACT FROM AUTHOR]

On 6 August 2012, the Curiosity rover successfully touched down on the Martian surface setting off the most ambitious surface exploration of this planetary body. Preceding this significant step were years of design, development, and testing of the Curiosity Entry, Descent, and Landing system to prepare for the most complex landing endeavor ever attempted at Mars. To address the numerous challenges, the approach and implementation of the overall Entry, Descent, and Landing verification and validation program relied on its decomposition into three distinct domains: flight dynamics, flight system and subsystem verification and validation. The test and analysis scope, the venues, and the processes utilized were tailored to each of these domains, and are discussed in greater detail. The overall lessons learned and conclusions described herein can serve as a pathfinder for the Entry, Descent, and Landing system testing approach and implementation of future Mars landed missions. [ABSTRACT FROM AUTHOR]

On 5 August 2012, the Mars Science Laboratory mission successfully landed Curiosity, the largest interplanetary rover ever built, on the surface of Mars. The entry, descent, and landing phase of this mission was by far the most complex landing ever attempted on a planetary body. The descent-stage propulsion subsystem was explicitly developed for the guided-entry and sky-crane maneuvers that enabled precise landing of Curiosity. Development of the descent-stage propulsion system required a number of new propulsion hardware developments, incorporating technologies not normally found in spacecraft propulsion subsystems. The in-flight performance of the descent-stage propulsion system was consistent with expectations, with the exception of higher than expected delivered thrust from the Mars lander engines. The descent-stage propulsion system completed its mission with significant capability margin with respect to thruster life, maximum delivered thrust, and available propellant at touchdown. [ABSTRACT FROM AUTHOR]

The Mars Science Laboratory is a scientific mission that placed a 900 kg rover with 10 science instruments on the surface of Mars. A key component for landing on the surface of Mars is the aeroshell, which decelerates the entry vehicle to just 1% of its initial kinetic energy at atmospheric entry. The forward portion of the aeroshell is the heatshield, which must be designed to withstand the pressure, heating, and shear stresses it experiences during the atmospheric deceleration process. The original thermal protection system chosen for the heatshield, SLA-561V, was the material used on all previous Mars aeroshelis but required new qualification testing. Unfortunately, the material failed testing in the more extreme environment three months after the aeroshell and project critical design reviews. A rapid decision was required to switch to another material under test by the Orion program called phenolic impregnated carbon ablator. With 18 months left to design, fabricate, test, and deliver this new heatshield for the 2009 launch opportunity, the heatshield team had to rapidly respond to the challenge and work in a manner that ensured success. Lessons learned can be had from both the failure of the original SLA-561V design and the success of the phenolic impregnated carbon ablator implementation. [ABSTRACT FROM AUTHOR]

On 5 August 2012, the Mars Science Laboratory entry vehicle successfully entered Mars's atmosphere and landed the Curiosity rover on its surface. The entry vehicle carried a system of heat shield instrumentation that measured the aerodynamic and aerothermal environment during entry. Included in these sensors were seven pressure transducers linked to ports across the entry vehicle forebody that recorded the local surface pressures during entry. These measured surface pressures were used to generate estimates of atmospheric quantities based on computationally modeled surface pressure distributions. Angle of attack, angle of sideslip, dynamic pressure, Mach number, atmospheric conditions, and vehicle aerodynamics were reconstructed from the measured pressures. Three data reduction algorithms using subsets of the available data were used to assess data quality and interpret the entry performance. A Kalman filter algorithm was used to combine all available data for a final, complete reconstruction. The reconstruction results indicate that the data quality was good, and aerodynamic performance was within the uncertainties of preflight models. The three independent reconstructions are in overall good agreement, with several small anomalies that were reconciled using reasonable corrections within known sources of error. The combined Kalman filter reconstruction identified winds that are reasonable compared with preflight atmospheric models. [ABSTRACT FROM AUTHOR]

The Mars Science Laboratory spacecraft landed an approximately 900 kg rover on Mars on 5 August 2012. Similar to past Mars missions, the spacecraft recorded inertial measurement unit data and radar altimeter measurements, but its aeroshell was also instrumented with flush atmospheric data system sensors that captured the pressure distribution on the vehicle during hypersonic and supersonic flight regimes. The rich data set enables a comprehensive postflight analysis of the vehicle's trajectory, atmosphere, and aerodynamics. This paper demonstrates a comprehensive statistical estimation methodology and applies it to the large but disparate data set to reconstruct the vehicle's entry performance, while using several statistical estimation methods, specifically the extended Kalman filter, unscented Kalman filter, and adaptive filter, which concurrently estimate both the state and the uncertainties. The results of the analysis shows good agreement between the estimated states, preflight predictions, and independent reconstructions, while also demonstrating the efficacy of the comprehensive statistical estimation method for Mars entry, descent, and landing reconstruction. [ABSTRACT FROM AUTHOR]

The Mars Science Laboratory, also called Curiosity, is a rover mission that launched on 26 November 2011 and landed on Mars 6 August 2012 in Gale Crater. The main challenges for the interplanetary navigation of the mission were to deliver the spacecraft to the correct interface point above the atmosphere of Mars and to accurately tell the spacecraft where it was as it entered the atmosphere. The spacecraft used guidance during its descent to Mars before the deployment of the parachute in order to minimize landing dispersions, resulting in a smaller landing zone that was closer to terrain of high scientific interest. This required an accurate delivery of the spacecraft to the entry interface and an update of the spacecraft state at entry, which was used to initialize the descent guidance system. Orbit determination during cruise was very accurate, being able to predict the line-of-sight position of the spacecraft after one week to within a few meters during the final weeks of approach. The spacecraft hit the top of the Martian atmosphere just 200 m from where it had been predicted to enter more than six days earlier, when it was still 2.6 million km away from Mars. This excellent level of accuracy was achieved by a combination of factors, including spacecraft characteristics, tracking data processing, dynamical modeling choices, and navigation filter setup. The accurate interplanetary navigation contributed to the very precise landing performance and to the overall success of the mission. [ABSTRACT FROM AUTHOR]

An entry, descent, and landing architecture capable of achieving Mars Science Laboratory-class landed accuracy (within 10 km of target) while delivering a Mars Exploration Rover-class payload to the surface of Mars is presented. The architecture consists of a Mars Exploration Rover-class aeroshell with a rigid, annular drag skirt. Maximum vehicle diameter, including drag skirt, is limited to be compatible with current launch-vehicle fairings. A single drag-skirt jettison event is used to control range during entry. Three-degree-of-freedom trajectory simulation is used in conjunction with Monte Carlo techniques to assess the flight performance of the proposed architecture. Results indicate that landed accuracy is competitive with preflight Mars Science Laboratory estimates, and peak heat rate and integrated heat load are significantly reduced relative to the Mars Exploration Rover entry system. Modeling parachute descent within the onboard guidance algorithm is found to remove range error bias present at touchdown; the addition of a range-based parachute deploy trigger is found to significantly improve landed accuracy. [ABSTRACT FROM AUTHOR]

The Mars Pathfinder probe entered the Martian atmosphere in 1997 and contained instrumentation that provided measurements of the superlight ablator (SLA) heat shield subsurface temperature at different locations during the entry sequence. These measurements represented the first Martian aeroheating flight data since the Viking Lander missions. The objective of this paper is to reconstruct the Pathfinder entry vehicle's aerothermal heating and heat shield material response using updated modeling tools and approaches in both direct and inverse manners. The direct approach consists of performing updated computational fluid dynamics calculations on a newly reconstructed entry trajectory to characterize the vehicle's heating environment. From the calculated heating boundary conditions, the heat shield in-depth temperature response is computed using an updated thermal response and ablation model for the SLA material. These predictions are compared directly to the flight data. In addition to the direct comparison approach, inverse methods are used to estimate boundary conditions that result in a closer match between the flight data and subsurface temperature predictions. The unblown surface heat transfer coefficient is reconstructed as a function of time using whole-time domain least-squares methods in conjunction with regularization techniques. [ABSTRACT FROM AUTHOR]

A tumbleweed rover is a spherical wind-driven vehicle designed to negotiate harsh Martian terrain while searching for areas of scientific interest. This paper models Martian rock fields and analyzes the rover's dynamics on these terrains. Computational procedures are established for creating randomized Martian rock fields based on statistical models. Optimization techniques allow for terrain generation to coincide with the rover's motion, reducing the computational cost of exploring broad regions. Efficient collision detection routines reduce the number of tests of potential collisions that must be performed between the rover and the terrain while establishing new contact constraints. Simulations demonstrate that bouncing is the rover's dominant mode of travel through the rock fields, and Monte Carlo simulations illustrate how the rover's downrange position depends on the rover design and atmospheric conditions. The rover's capacity for long-distance travel over Martian rock fields is verified. [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]

Spectroscopic observations of the 2006 Stardust Sample Return Capsule entry are presented, obtained by means of a slitless miniature echelle spectrograph onboard NASA's DC-8 airborne laboratory. The data cover the wavelength range from 336 to 880 nm, at 0.14-0.9 nm resolution, and were obtained during the time interval when radiative heating was most important. The data contain a broadband continuum, presumably from the hot heat-shield surface, shock-layer air plasma emissions of N, O, and N2+, and atomic hydrogen and CN molecular band emission from the ablating heat-shield material, a form of phenol-impregnated carbon ablator. Early in flight, there were also atomic line emissions of Zn, K, Ca+, and Na, presumably from a white Z-93P paint applied to the top of the phenolimpregnated carbon ablator. At each moment along the trajectory, the whole spectrum was recorded simultaneously, but broken into smaller segments. Key issues addressed in the data reduction and calibration are described. The interpretation of these data was given elsewhere. [ABSTRACT FROM AUTHOR]

In the framework of Research Task Group 043 of the NATO Research and Technology Organization, an analysis of the capabilities in prediction of aerothermal loads acting on a crew exploration vehicle at the higher altitudes of its reentry trajectory has been performed. In particular, the focus of this investigation is to provide information where overlap between the continuum approach (i.e., computational fluid dynamics) and particle approach (i.e., direct simulation Monte Carlo) occurs and improves understanding of relevant physics in transitional regime. Computational fluid dynamics calculations with slip flow boundary conditions had shown good predicting capabilities of slip velocity, slip temperature, and pressure, but not the surface heat flux. The analysis of the contributions to the total surface heat flux had exhibited the lack of continuum model in evaluating the convective heat flux in the transitional zone of the Orion reentry trajectory and an underestimation of shock wave thickness. [ABSTRACT FROM AUTHOR]

An investigation of the aeroheating environment of the Project Orion Crew Exploration Vehicle was performed in the Arnold Engineering Development Center Hypervelocity Wind Tunnel 9 Mach 8 and Mach 10 nozzles and in the NASA Langley Research Center 20-Inch Mach 6 Air Tunnel. Heating data were obtained using a thermocouple- instrumented ~ 0.035-scale model [0.1778 m (7 in.) diameter] of the flight vehicle. Runs were performed in the Tunnel 9 Mach 10 nozzle at freestream unit Reynolds numbers of 1 x 106 to 20 x 106/f t, in the Tunnel 9 Mach 8 nozzle at freestream unit Reynolds numbers of 8 x 106 to 48 x 106/f t, and in the 20-Inch Mach 6 Air Tunnel at freestream unit Reynolds numbers of 1 x 106 to 7 x 106/ft. In both facilities, enthalpy levels were low and the test gas (N2 in Tunnel 9 and air in the 20-Inch Mach 6 Air Tunnel) behaved as a perfect gas. These test conditions produced laminar, transitional, and turbulent data in the Tunnel 9 Mach 10 nozzle; transitional and turbulent data in the Tunnel 9 Mach 8 nozzle; and laminar and transitional data in the 20-Inch Mach 6 Air Tunnel. Laminar and turbulent predictions were generated for all wind-tunnel test conditions, and comparisons were performed with the experimental data to help define the accuracy of the computational method. In general, it was found that both laminar data and predictions and turbulent data and predictions agreed to within less than the estimated ± 12% experimental uncertainty estimate. Laminar heating distributions from all three data sets were shown to correlate well and demonstrated Reynolds numbers independence when expressed in terms of the Stanton number based on adiabatic-wall-recovery enthalpy. Transition-onset locations on the lee-side centerline were determined from the data and correlated in terms of boundary-layer parameters. Finally, turbulent heating augmentation ratios were determined for several body-point locations and correlated in terms of the boundary-layer momentum Reynolds number. [ABSTRACT FROM AUTHOR]

An experimental wind-tunnel program is being conducted in support of a NASA wide effort to develop a Space Shuttle replacement and to support the agency's long-term objective of returning to the moon and Mars. This paper documents experimental measurements made on several scaled ceramic heat transfer models of the proposed Project Orion Crew Exploration Vehicle. The experimental data highlighted in this paper will be used to assess numerical tools that will be used to generate the flight aerothermodynamic database. Global heat transfer images and centerline heat transfer distributions were obtained over a range of freestream Reynolds numbers and angles of attack with the phosphor thermography technique. Temperature data were measured on the forebody and afterbody and were used to compute the heating on the vehicle as well as the boundary-layer state on the forebody surface. Several model support configurations were assessed to minimize potential aftbody support interference. Although naturally fully developed turbulent levels were not obtained on the forebody, the use of boundary-layer trips generated fully developed turbulent flow. Laminar and turbulent computational results are shown to be in good agreement with the data. In addition, the ability of the global phosphor thermography method to provide quantitative heating measurements in the low-temperature environment of the capsule base region was assessed and the lack of significant model support hardware influence on heating was shown. [ABSTRACT FROM AUTHOR]

The Crew Exploration Vehicle presently being developed to return humans from lunar missions will typically experience entry environments far more severe than those for return from low-Earth orbit. Certification of the thermal protection system materials for these environments, and ultimately for Mars return, will require extensive testing. As part of an effort to bound the required testing capability, trajectories and the associated aeroheating environments were generated for more than 60 unique entry cases. Using the Apollo Command Module as the baseline entry system, trajectories for a range of lunar and Mars return Earth-entry scenarios were developed using 3-degree of freedom trajectory simulation. For direct entry, a matrix of cases that reflects expected minimum and maximum values of vehicle ballistic coefficient, inertial velocity, and flight-path angle at entry interface was considered. For aerocapture, a range of values of initial velocity and ballistic coefficient was examined that, when combined with appropriate initial flight-path angles, enable the vehicle to achieve low-Earth orbit by employing either a full-lift-vector-up or full-lift-vector-down attitude. For each trajectory, aeroheating environments, intended to bound the thermal protection material system requirements for likely concepts, were generated using engineering methods. The trades examined in this study distinguished the classes of missions/concepts that will require ablative systems as well as those for which reusable systems may be feasible. Results highlight those entry conditions and modes suitable for human flight, considering vehicle deceleration levels experienced during entry. The aeroheating environments generatedin this study can be used by the vehicle designer to assess the material testing limits and facility requirements for a broad range of concepts and missions. [ABSTRACT FROM AUTHOR]

Computational simulations and analyses may be used to quantify the risk envelope for crew escape resulting from an accidental detonation of explosive materials at a specified time during the ascent phase for crew exploration vehicles. Here, a methodology was developed and applied to assess the explosive risk potential, focusing on internal detonations or explosions. The approach was to computationally model the geometry of a launch vehicle and the time evolution of an accident scenario with explosive detonation. A series of detailed detonation simulations were completed for three different crew exploration vehicle launch vehicle configurations. All simulations assumed the formation of an explosive, mixture of liquid oxygen and liquid hydrogen. Two different energy-release models were used in this study. One was based on programmed burn with a Jones-Wilkins-Lee analytical description of the reaction-product expansion isentrope, and the other used a volumetric energy source formulation. Based on this methodology and simulations, the predicted pressure field had significant magnitudes in the region of the service module and crew exploration vehicle. In fact, the overpressure is so severe that it would not be a survivable event for the crew. [ABSTRACT FROM AUTHOR]

In July 2006, the NASA Administrator's Office and the Constellation Program Office jointly commissioned a propellant options study to perform a comprehensive detailed analysis of the major propulsion system alternatives for the planned crew exploration vehicle and lunar surface access module before the crew exploration vehicle systems requirements review. The study team, which involved over 50 NASA engineers and analysts, performed detailed modeling of the propulsion systems and the crew exploration vehicle and lunar surface access module configurations to calculate the integrated system-level impacts on payload, mass,*reliability, safety, risk, and life-cycle costs. This paper presents an overview of the tools and methods used to assess the propulsion system and architecture-level effects of employing the propellant options. It also provides the architecture-level relative performance, life-cycle costs, reliability, safety, schedule, programmatic risk, and sustainability assessments for scenarios that employ a wide variety of main and auxiliary propellant combinations. It also includes a qualitative discussion of the major risks of each scenario. As a result of the study, pressure-fed hypergolic propellants were selected as the baseline for the crew exploration vehicle service module main and auxiliary propulsion systems. Pump-fed hydrogen propellants were kept on the descent stage of the lunar surface access module, but the lunar surface access module ascent stage propellant baseline was changed to pressure-fed methane for the main and auxiliary propulsion systems, with a recommendation that the trade space be kept open while more detailed configuration studies are performed. [ABSTRACT FROM AUTHOR]

A Mars Tumbleweed rover is a spherical, wind-driven, planetary rover. Compared with conventional rovers, a tumbleweed rover can travel farther faster and gain access to areas such as valleys and chasms that previously were inaccessible. This paper presents design, mathematical modeling, computer simulation, and testing of various tumbleweed rover concepts. In particular, we present wind-tunnel data indicating that a box-kite configuration represents a promising tumbleweed rover design, we show that a working box-kite-type tumbleweed can be constructed, and we show that center of mass variation shows promise as the basis of a tumbleweed rover navigation system. [ABSTRACT FROM AUTHOR]

The impending development of NASA's Orion crew exploration vehicle will require a new entry guidance algorithm that provides sufficient performance to meet all requirements. This study examined the effects on entry footprints of enhancing the skip trajectory entry guidance used in the Apollo program. The skip trajectory entry guidance was modified to include a numerical predictor-corrector phase during the atmospheric skip portion of the entry trajectory. A 4-degree-of-freedom simulation was used to determine the range capability of the entry vehicle for the baseline Apollo entry guidance and the predictor-corrector enhanced guidance with both high and low lofting at several lunar return entry conditions. The results show that the predictor-corrector guidance modification significantly improves the entry range capability of the crew exploration vehicle for the lunar return mission. The performance provided by the enhanced algorithm is likely to meet the entry range requirements for the crew exploration vehicle. [ABSTRACT FROM AUTHOR]

The article provides information on the Mars Science Laboratory (MSL) entry, descent and landing (EDL) sequence, which is designed to place a rover on the Mars surface at near-zero velocity.