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NASA NTRS2002

A Computational Intelligence (CI) Approach to the Precision Mars Lander Problem

A Mars precision landing requires a landed footprint of no more than 100 meters. Obstacles to reducing the landed footprint include trajectory dispersions due to initial atmospheric entry conditions such as entry angle, parachute deployment height, environment parameters such as wind, atmospheric density, parachute deployment dynamics, unavoidable injection error or propagated error from launch, etc. Computational Intelligence (CI) techniques such as Artificial Neural Nets and Particle Swarm Optimization have been shown to have great success with other control problems. The research period extended previous work on investigating applicability of the computational intelligent approaches. The focus of this investigation was on Particle Swarm Optimization and basic Neural Net architectures. The research investigating these issues was performed for the grant cycle from 5/15/01 to 5/15/02. Matlab 5.1 and 6.0 along with NASA's POST were the primary computational tools.

Birge, Brian↗
NASA NTRS2002

Mars-Gram Validation with Mars Global Surveyor Data

Mars Global Reference Atmospheric Model (Mars-GRAM 2001) is an engineering-level Mars atmosphere model widely used for many b4ars mission applications. From 0-80 km, it is based on NASA Ames Mars General Circulation Model (MGCM), while above 80 km it is based on University of Arizona Mars Thermospheric General Circulation Model. Mars-GRAM 2001 and MGCM use surface topography from Mars Global Surveyor Mars Orbiting Laser Altimeter (MOLA). Validation studies are described comparing Mars-GRAM with Mars Global Surveyor Radio Science (RS) and Thermal Emission Spectrometer (TES) data. RS data from 2480 profiles were used, covering latitudes 75deg S to 72deg N, surface to approx. 40 km, for seasons ranging from areocentric longitude of Sun (Ls) = 70-160deg and 265-310deg. RS data spanned a range of local times, mostly 0-9 hours and 18-24 hours. For interests in aerocapture and precision landing, comparisons concentrated on atmospheric density. At a fixed height of 20 km, measured RS density varied by about a factor of 2.5 over the range of latitudes and Ls values observed. Evaluated at matching positions and times, average RS/Mars-GRAM density ratios were generally lf0.05, except at heights above approx. 25 km and latitudes above approx.50deg N. Average standard deviation of RS/Mars-GRAM density ratio was 6%. TES data were used covering surface to approx. 40 km, over more than a full Mars year (February, 1999 - June, 2001, just before start of Mars global dust storm). Depending on season, TES data covered latitudes 85deg S to 85deg N. Most TES data were concentrated near local times 2 hours and 14 hours. Observed average TES/Mars-GRAM density ratios were generally 1+/-0.05, except at high altitudes (15-30 km, depending on season) and high latitudes (> 45deg N), or at most altitudes in the southern hemisphere at Ls approx. 90 and 180deg). Compared to TES averages for a given latitude and season, TES data had average density standard deviation about the mean of approx. 6.5-10.5% (varying with height) for all data, or approx. 5- 12%, depending on time of day and dust optical depth. Average standard deviation of TES/Mars-GRAM density ratio was 8.9% for local time 2 hours and 7.1% for 1o:al time 14 hours. Thus standard deviation of observed TES/Mars-GRAM density ratio, evaluated at matching positions and times, is about the same as the standard deviation of TES data about the TES mean value at a given position and season.

Justus, C. G.↗
NASA NTRS2002

Experimental Hypersonic Aerodynamic Characteristics of the 2001 Mars Surveyor 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 NASA Langley Aerothermodynamic Laboratory for post-normal shock Reynolds numbers (based on forebody diameter) ranging from 2,637 to 92,350, angles of attack ranging from 0 tip to 23 degrees at 0 and 2 degree sideslip, and normal-shock density ratios of 5 and 12. Based upon the proposed entry trajectory of the 2001 Lander, the blunt body heavy gas tests in CF, 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 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 were 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 L/D for precision landing. Inviscid hypersonic flow computations using an unstructured grid were made to provide a quick assessment of the Lander aerodynamics. Navier-Stokes computational predictions were found to be in very good agreement with experimental measurement.

Horvath, Thomas J.↗
NASA NTRS2002

Control Surface and Afterbody Experimental Aeroheating for a Proposed Mars Smart Lander Aeroshell

Several configurations, having a Viking aeroshell heritage and providing lift-to-drag required for precision landing, have been considered for a proposed Mars Smart Lander. An experimental aeroheating investigation of two configurations, one having a blended tab and the other a blended shelf control surface, has been conducted at the NASA Langley Research Center in the 20-Inch Mach 6 Air Tunnel to assess heating levels on these control surfaces and their effects on afterbody heating. The proposed Mars Smart Lander concept is to be attached through its aeroshell to the main spacecraft bus, thereby producing cavities in the forebody heat shield upon separation prior to entry into the Martian atmosphere. The effects these cavities will have on the heating levels experienced by the control surface and the afterbody were also examined. The effects of Reynolds number, angle-of-attack, and cavity location on aeroheating levels and distributions were determined and are presented. At the highest angle-of-attack, blended tab heating was increased due to transitional reattachment of the separated shear layer. The placement of cavities downstream of the control surface greatly influenced aeroheating levels and distributions. Forebody heat shield cavities had no effect on afterbody heating and the presence of control surfaces decreased leeward afterbody heating slightly.

Liechty, Derek S.↗
NASA NTRS2003

Computational Aeroheating Predictions for Mars Lander Configurations

The proposed Mars Science Laboratory (MSL) mission is intended to deliver a large rover to the Martian surface within 10 km of the target site. This paper presents computational fluid dynamics (CFD) predictions of forebody heating rates for two MSL entry configurations with fixed aerodynamic trim tabs. Results are compared to heating on a 70-deg sphere-cone reference geometry. All three heatshield geometries are designed to trim hypersonically at a 16 deg angle of attack in order to generate the lift-to-drag ratio (LID) required for precision landing. Comparisons between CFD and tunnel data are generally in good agreement for each configuration, but the computations predict more flow separation and higher heating on a trim tab inclined 10 deg relative to the surface. CFD solutions at flight conditions were obtained using an 8-species Mars gas in chemical and thermal non-equilibrium. Laminar and Baldwin-Lomax solutions were used to estimate the effects of the trim tabs and turbulence on heating. A tab extending smoothly from the heatshield flank is not predicted to increase laminar or turbulent heating rates above the reference levels. Laminar heating on a tab deflected 10 deg from the conical heatshield is influenced by flow separation and is up to 35% above the baseline heating rate. The turbulent solution on the inclined tab configuration predicts attached flow and a 43% heating increase above the reference level.

Edquist, Karl T.↗
NASA NTRS2003

Mars Global Reference Atmospheric Model (Mars-GRAM) and Database for Mission Design

Mars Global Reference Atmospheric Model (Mars-GRAM 2001) is an engineering-level Mars atmosphere model widely used for many Mars mission applications. From 0-80 km, it is based on NASA Ames Mars General Circulation Model, while above 80 km it is based on Mars Thermospheric General Circulation Model. Mars-GRAM 2001 and MGCM use surface topography from Mars Global Surveyor Mars Orbiting Laser Altimeter. Validation studies are described comparing Mars-GRAM with Mars Global Surveyor Radio Science and Thermal Emission Spectrometer data. RS data from 2480 profiles were used, covering latitudes 75 deg S to 72 deg N, surface to approximately 40 km, for seasons ranging from areocentric longitude of Sun (Ls) = 70-160 deg and 265-310 deg. RS data spanned a range of local times, mostly 0-9 hours and 18-24 hours. For interests in aerocapture and precision landing, comparisons concentrated on atmospheric density. At a fixed height of 20 km, RS density varied by about a factor of 2.5 over ranges of latitudes and Ls values observed. Evaluated at matching positions and times, these figures show average RSMars-GRAM density ratios were generally 1+/-)0.05, except at heights above approximately 25 km and latitudes above approximately 50 deg N. Average standard deviation of RSMars-GRAM density ratio was 6%. TES data were used covering surface to approximately 40 km, over more than a full Mars year (February, 1999 - June, 2001, just before start of a Mars global dust storm). Depending on season, TES data covered latitudes 85 deg S to 85 deg N. Most TES data were concentrated near local times 2 hours and 14 hours. Observed average TES/Mars-GRAM density ratios were generally 1+/-0.05, except at high altitudes (15-30 km, depending on season) and high latitudes (greater than 45 deg N), or at most altitudes in the southern hemisphere at Ls approximately 90 and 180 deg. Compared to TES averages for a given latitude and season, TES data had average density standard deviation about the mean of approximately 2.5% for all data, or approximately 1-4%, depending on time of day and dust optical depth. Average standard deviation of TES/Mars-GRAM density ratio was 8.9% for local time 2 hours and 7.1% for local time 14 hours. Thus standard deviation of observed TES/Mars-GRAM density ratio, evaluated at matching positions and times, is about three times the standard deviation of TES data about the TES mean value at a given position and season.

Justus, C. G.↗
NASA NTRS2003

Computational Aeroheating Predictions for Mars Lander Configurations

The proposed Mars Science Laboratory (MSL) mission is intended to deliver a large rover to the Martian surface within 10 km of the target site. This paper presents computational fluid dynamics (CFD) predictions of forebody heating rates for two MSL entry configurations with fixed aerodynamic trim tabs. Results are compared to heating on a 70-deg sphere-cone reference geometry. All three heatshield geometries are designed to trim hypersonically at a 16 deg angle of attack in order to generate the lift-to-drag ratio (L/D) required for precision landing. Comparisons between CFD and tunnel data are generally in good agreement for each configuration, but the computations predict more flow separation and higher heating on a trim tab inclined 10 deg relative to the surface. CFD solutions at flight conditions were obtained using an 8-species Mars gas in chemical and thermal nonequilibrium. Laminar and Baldwin-Lomax solutions were used to estimate the effects of the trim tabs and turbulence on heating. A tab extending smoothly from the heatshield flank is not predicted to increase laminar or turbulent heating rates above the reference levels. Laminar heating on a tab deflected 10 deg from the conical heatshield is influenced by flow separation and is up to 35% above the baseline heating rate. The turbulent solution on the inclined tab configuration predicts attached flow and a 43% heating increase above the reference level.

Edquist, Karl T.↗
NASA NTRS1998

An Atmospheric Guidance Algorithm Testbed for the Mars Surveyor Program 2001 Orbiter and Lander

An Atmospheric Flight Team was formed by the Mars Surveyor Program '01 mission office to develop aerocapture and precision landing testbed simulations and candidate guidance algorithms. Three- and six-degree-of-freedom Mars atmospheric flight simulations have been developed for testing, evaluation, and analysis of candidate guidance algorithms for the Mars Surveyor Program 2001 Orbiter and Lander. These simulations are built around the Program to Optimize Simulated Trajectories. Subroutines were supplied by Atmospheric Flight Team members for modeling the Mars atmosphere, spacecraft control system, aeroshell aerodynamic characteristics, and other Mars 2001 mission specific models. This paper describes these models and their perturbations applied during Monte Carlo analyses to develop, test, and characterize candidate guidance algorithms.

Striepe, Scott A.↗
NASA NTRS2005

GLOBAL REFERENCE ATMOSPHERIC MODELS FOR AEROASSIST APPLICATIONS

Aeroassist is a broad category of advanced transportation technology encompassing aerocapture, aerobraking, aeroentry, precision landing, hazard detection and avoidance, and aerogravity assist. The eight destinations in the Solar System with sufficient atmosphere to enable aeroassist technology are Venus, Earth, Mars, Jupiter, Saturn, Uranus, Neptune, and Saturn's moon Titan. Engineering-level atmospheric models for five of these targets - Earth, Mars, Titan, Neptune, and Venus - have been developed at NASA's Marshall Space Flight Center. These models are useful as tools in mission planning and systems analysis studies associated with aeroassist applications. The series of models is collectively named the Global Reference Atmospheric Model or GRAM series. An important capability of all the models in the GRAM series is their ability to simulate quasi-random perturbations for Monte Carlo analysis in developing guidance, navigation and control algorithms, for aerothermal design, and for other applications sensitive to atmospheric variability. Recent example applications are discussed.

Duvall, Aleta↗
NASA NTRS2002

An overview of the aerocapture flight test experiment (AFTE)

The paper describes the technical details underlying the AFTE mission scenario and discusses how the results would benefit the future aerocapture missions and related hypersonic flight applications like precision landing.

aerocapture aeroshell flight test↗
NASA NTRS2006

Telecommunications IT and navigation for future Mars exploration missions 2006 IEEE Aerospace Conference

There are three primary drivers behind current investments in telecommunications information technology and navigation. One is finding ways to maximize the volume of science data returned from missions since i nstrument data generation often exceeds communication bandwidth. Another is to provide the necessary technology to enable networked spacecraft around Mars. The third driver is to enable more precise landing so in-situ vehicles can be placed in the more scientifically interesting regions. This paper describes the current NASA investments in these areas funded through the NASA Mars Technology Base Program NRA.

data compression↗
NASA NTRS2006

The JPL roadmap for Deep Space navigation

This paper reviews the tentative set of deep space missions that will be supported by NASA's Deep Space Mission System in the next twenty-five years, and extracts the driving set of navigation capabilities that these missions will require. There will be many challenges including the support of new mission navigation approaches such as formation flying and rendezvous in deep space, low-energy and low-thrust orbit transfers, precise landing and ascent vehicles, and autonomous navigation. Innovative strategies and approaches will be needed to develop and field advanced navigation capabilities.

navigation↗
NASA NTRS2007

Coherent Lidar Activities at NASA Langley Research Center

NASA Langley Research Center has been developing and using coherent lidar systems for many years. The current projects at LaRC are the Global Wind Observing Sounder (GWOS) mission preparation, the Laser Risk Reduction Program (LRRP), the Instrument Incubator Program (IIP) compact, rugged Doppler wind lidar project, the Autonomous precision Landing and Hazard detection and Avoidance Technology (ALHAT) project for lunar landing, and the Skywalker project to find and use thermals to extend UAV flight time. These five projects encompass coherent lidar technology development; characterization, validation, and calibration facilities; compact, rugged packaging; computer simulation; trade studies; data acquisition, processing, and display development; system demonstration; and space mission design. This paper will further discuss these activities at LaRC.

Kavaya, Michael J.↗
NASA NTRS2006

Overview of a Proposed Flight Validation of Aerocapture System Technology

Aerocapture is a very useful capability for NASA that can be used across a wide range of planetary mission sizes and destinations. A substantial mass advantage may be realized through aerocapture maneuver implementation. The mass advantage is enabling for certain outer planet mission profiles. Aerocapture technology provides corollary benefits to the related applications of atmospheric entry and precision landing on worlds with atmospheres through aero/aerothermodynamic model validation, hypersonic guided flight, tps materials, and performance model validation. The ST9 Aerocapture flight validation will be sufficient to immediately infuse aerocapture technology into future NASA science missions. The advanced technologies being flight validated will enable the system level goal of performing an aerocapture maneuver. The advanced technologies include: The GN&C System, TPS materials, plus Advanced recession and heat flux sensors.

Keys, Andrew S.↗
NASA NTRS2009

Apollo Mission Techniques Lunar Orbit Activities - Part 1a

This slide presentation reviews the planned sequence of events and the rationale for all lunar missions, and the flight experiences and lessons learned for the lunar orbit activities from a trajectory perspective. Shown are trajectories which include the moon's position at the various stages in the complete trip from launch, to the return and reentry. Included in the presentation are objectives and the sequence of events,for the Apollo 8, and Apollo 10. This is followed by a discussion of Apollo 11, including: the primary mission objective, the sequence of events, and the flight experience. The next mission discussed was Apollo 12. It reviews the objectives, the ground tracking, procedure changes, and the sequence of events. The aborted Apollo 13 mission is reviewed, including the objectives, and the sequence of events. Brief summaries of the flight experiences for Apollo 14-16 are reviewed. The flight sequence of events of Apollo 17 are discussed. In summary each mission consistently performing precision landings required that Apollo lunar orbit activities devote considerable attention to: (1) Improving fidelity of lunar gravity models, (2) Maximizing availability of ground tracking, (3) Minimizing perturbations on the trajectory, (4) Maximizing LM propellant reserves for hover time. Also the use of radial separation maneuvers (1) allows passive re-rendezvous after each rev, but ... (2) sensitive to small dispersions in initial sep direction

Interbartolo, Michael A.↗
NASA NTRS2009

Transition Analysis for the Mars Science Laboratory Entry Vehicle

Viscous Laminar-turbulent transition plays an important role in the design of the Mars Science Laboratory (MSL) entry vehicle. The lift-to-drag ratio required for the precision landing trajectory will be achieved via an angle of attack equal to 16 degrees. At this relatively high angle of attack, the boundary layer flow near the leeward meridian is expected to transition early in the trajectory, resulting in substantially increased heating loads. This paper presents stability calculations and transition correlations for a series of wind tunnel models of the MSL vehicle. Experimentally measured transition onset locations are used to correlate with the N-factor calculations for various wind tunnel conditions. Due to relatively low post-shock Mach numbers near the edge of the boundary layer, the dominant instability waves are found to be of the first mode type. The N-factor values correlating with measured transition onset at selected test points from the Mach 6 conventional facility experiments fall between 3.5 and 4.5 and apparently vary linearly with the wind tunnel unit Reynolds number, indicating strong receptivity effect. The small transition N value is consistent with previous correlations for second-mode dominant transition in the same wind tunnel facility. Stability calculations for stationary and traveling crossflow instability waves in selected configurations indicate that an N value of 4 and 6, respectively, correlates reasonably well with transition onset discerned from one experimentally measured thermographic image.

Chang, Chau-Lyan↗
NASA NTRS2010

Present Challenges, Critical Needs, and Future Technological Directions for NASA's GN and C Engineering Discipline

The National Aeronautics and Space Administration (NASA) is currently undergoing a substantial redirection. Notable among the changes occurring within NASA is the stated emphasis on technology development, integration, and demonstration. These new changes within the Agency should have a positive impact on the GN&C discipline given the potential for sizeable investments for technology development and in-space demonstrations of both Autonomous Rendezvous & Docking (AR&D) systems and Autonomous Precision Landing (APL) systems. In this paper the NASA Technical Fellow for Guidance, Navigation and Control (GN&C) provides a summary of the present technical challenges, critical needs, and future technological directions for NASA s GN&C engineering discipline. A brief overview of the changes occurring within NASA that are driving a renewed emphasis on technology development will be presented as background. The potential benefits of the planned GN&C technology developments will be highlighted. This paper will provide a GN&C State-of-the-Discipline assessment. The discipline s readiness to support the goals & objectives of each of the four NASA Mission Directorates is evaluated and the technical challenges and barriers currently faced by the discipline are summarized. This paper will also discuss the need for sustained investments to sufficiently mature the several classes of GN&C technologies required to implement NASA crewed exploration and robotic science missions.

Dennehy, Cornelius J.↗