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NASA NTRS2005

A Proposed Role of Aeroelasticity in NASA's New Exploration Vision

On 14 January 2004, NASA received a mandate to return astronauts to the Moon, evolve a sustained presence there, then head out into the solar system to Mars and perhaps beyond. This new space exploration initiative directs NASA to develop human and robotic technologies that can deliver payloads larger than Apollo to the Moon, to Mars, and bring astronauts and samples safely back to Earth at costs much lower than Apollo. These challenges require creative aerospace systems. On proposed technology for safely delivering payloads to the surface of Mars and returning samples to Earth involves deployed flexible and inflatable decelerators for atmospheric entry. Because inflatable decelerators provide the entry vehicle more drag surface area at smaller mass than traditional ablative devices, this class of decelerators can potentially accomodate larger mass payloads. The flexibility of these lightweight aeroshells can pose both vehicle and aeroelastic stability problems if not properly designed for the expected flight regimes. Computational tools need to be developed for modelling the large and nonlinear deformations of these highly flexible structures. Unlike wind tunnel testing, an integrated and efficient aeroelastic analysis tool can explore the entire flight environment. This paper will provide some background on flexible deployable decelerators, survey the current state of technology and outline the proposed development of an aeroelastic analysis and capability.

Bartels, Robert E.↗

NASA NTRS2005

Aerocapture Performance Analysis of A Venus Exploration Mission

A performance analysis of a Discovery Class Venus Exploration Mission in which aerocapture is used to capture a spacecraft into a 300km polar orbit for a two year science mission has been conducted to quantify its performance. A preliminary performance assessment determined that a high heritage 70 sphere-cone rigid aeroshell with a 0.25 lift to drag ratio has adequate control authority to provide an entry flight path angle corridor large enough for the mission s aerocapture maneuver. A 114 kilograms per square meter ballistic coefficient reference vehicle was developed from the science requirements and the preliminary assessment s heating indicators and deceleration loads. Performance analyses were conducted for the reference vehicle and for sensitivity studies on vehicle ballistic coefficient and maximum bank rate. The performance analyses used a high fidelity flight simulation within a Monte Carlo executive to define the aerocapture heating environment and deceleration loads and to determine mission success statistics. The simulation utilized the Program to Optimize Simulated Trajectories (POST) that was modified to include Venus specific atmospheric and planet models, aerodynamic characteristics, and interplanetary trajectory models. In addition to Venus specific models, an autonomous guidance system, HYPAS, and a pseudo flight controller were incorporated in the simulation. The Monte Carlo analyses incorporated a reference set of approach trajectory delivery errors, aerodynamic uncertainties, and atmospheric density variations. The reference performance analysis determined the reference vehicle achieves 100% successful capture and has a 99.87% probability of attaining the science orbit with a 90 meters per second delta V budget for post aerocapture orbital adjustments. A ballistic coefficient trade study conducted with reference uncertainties determined that the 0.25 L/D vehicle can achieve 100% successful capture with a ballistic coefficient of 228 kilograms per square meter and that the increased ballistic coefficient increases post aerocapture V budget to 134 meters per second for a 99.87% probability of attaining the science orbit. A trade study on vehicle bank rate determined that the 0.25 L/D vehicle can achieve 100% successful capture when the maximum bank rate is decreased from 30 deg/s to 20 deg/s. The decreased bank rate increases post aerocapture delta V budget to 102 meters per second for a 99.87% probability of attaining the science orbit.

Starr, Brett R.↗

NASA NTRS2006

AEROELASTIC SIMULATION TOOL FOR INFLATABLE BALLUTE AEROCAPTURE

A multidisciplinary analysis tool is under development for predicting the impact of aeroelastic effects on the functionality of inflatable ballute aeroassist vehicles in both the continuum and rarefied flow regimes. High-fidelity modules for continuum and rarefied aerodynamics, structural dynamics, heat transfer, and computational grid deformation are coupled in an integrated multi-physics, multi-disciplinary computing environment. This flexible and extensible approach allows the integration of state-of-the-art, stand-alone NASA and industry leading continuum and rarefied flow solvers and structural analysis codes into a computing environment in which the modules can run concurrently with synchronized data transfer. Coupled fluid-structure continuum flow demonstrations were conducted on a clamped ballute configuration. The feasibility of implementing a DSMC flow solver in the simulation framework was demonstrated, and loosely coupled rarefied flow aeroelastic demonstrations were performed. A NASA and industry technology survey identified CFD, DSMC and structural analysis codes capable of modeling non-linear shape and material response of thin-film inflated aeroshells. The simulation technology will find direct and immediate applications with NASA and industry in ongoing aerocapture technology development programs.

Liever, P. A.↗

NASA NTRS2005

Mars Aerocapture and Validation of Mars-GRAM with TES Data

Mars Global Reference Atmospheric Model (Mars-GRAM) is a widely-used engineering- level Mars atmospheric model. Applications include systems design, performance analysis, and operations planning for aerobraking, entry descent and landing, and aerocapture. Typical Mars aerocapture periapsis altitudes (for systems with rigid-aeroshell heat shields) are about 50 km. This altitude is above the 0-40 km height range covered by Mars Global Surveyor Thermal Emission Spectrometer (TES) nadir observations. Recently, TES limb sounding data have been made available, spanning more than two Mars years (more than 200,000 data profiles) with altitude coverage up to about 60 km, well within the height range of interest for aerocapture. Results are presented comparing Mars-GRAM atmospheric density with densities from TES nadir and limb sounding observations. A new Mars-GRAM feature is described which allows individual TES nadir or limb profiles to be extracted from the large TES databases, and to be used as an optional replacement for standard Mars-GRAM background (climatology) conditions. For Monte-Carlo applications such as aerocapture guidance and control studies, Mars-GRAM perturbations are available using these TES profile background conditions.

Justus, C. G.↗

NASA NTRS2006

Aerocapture Systems Analysis for a Titan Mission

Performance projections for aerocapture show a vehicle mass savings of between 40 and 80%, dependent on destination, for an aerocapture vehicle compared to an all-propulsive chemical vehicle. In addition aerocapture is applicable to multiple planetary exploration destinations of interest to NASA. The 2001 NASA In-Space Propulsion Program (ISP) technology prioritization effort identified aerocapture as one of the top three propulsion technologies for solar system exploration missions. An additional finding was that aerocapture needed a better system definition and that supporting technology gaps needed to be identified. Consequently, the ISP program sponsored an aerocapture systems analysis effort that was completed in 2002. The focus of the effort was on aerocapture at Titan with a rigid aeroshell system. Titan was selected as the initial destination for the study due to potential interest in a follow-on mission to Cassini/Huygens. Aerocapture is feasible, and the performance is adequate, for the Titan mission and it can deliver 2.4 times more mass to Titan than an all-propulsive system for the same launch vehicle.

Lockwood, Mary K.↗

NASA NTRSDate not supplied

Aerodynamic Decelerators for Planetary Exploration: Past, Present, and Future

In this paper, aerodynamic decelerators are defined as textile devices intended to be deployed at Mach numbers below five. Such aerodynamic decelerators include parachutes and inflatable aerodynamic decelerators (often known as ballutes). Aerodynamic decelerators play a key role in the Entry, Descent, and Landing (EDL) of planetary exploration vehicles. Among the functions performed by aerodynamic decelerators for such vehicles are deceleration (often from supersonic to subsonic speeds), minimization of descent rate, providing specific descent rates (so that scientific measurements can be obtained), providing stability (drogue function - either to prevent aeroshell tumbling or to meet instrumentation requirements), effecting further aerodynamic decelerator system deployment (pilot function), providing differences in ballistic coefficients of components to enable separation events, and providing height and timeline to allow for completion of the EDL sequence. Challenging aspects in the development of aerodynamic decelerators for planetary exploration missions include: deployment in the unusual combination of high Mach numbers and low dynamic pressures, deployment in the wake behind a blunt-body entry vehicle, stringent mass and volume constraints, and the requirement for high drag and stability. Furthermore, these aerodynamic decelerators must be qualified for flight without access to the exotic operating environment where they are expected to operate. This paper is an introduction to the development and application of aerodynamic decelerators for robotic planetary exploration missions (including Earth sample return missions) from the earliest work in the 1960s to new ideas and technologies with possible application to future missions. An extensive list of references is provided for additional study.

Juan R Cruz↗

NASA NTRS2006

Mars Aerocapture Systems Study

Mars Aerocapture Systems Study (MASS) is a detailed study of the application of aerocapture to a large Mars robotic orbiter to assess and identify key technology gaps. This study addressed use of an Opposition class return segment for use in the Mars Sample Return architecture. Study addressed mission architecture issues as well as system design. Key trade studies focused on design of aerocapture aeroshell, spacecraft design and packaging, guidance, navigation and control with simulation, computational fluid dynamics, and thermal protection system sizing. Detailed master equipment lists are included as well as a cursory cost assessment.

Wright, Henry S.↗

NASA NTRS2005

An Investigation of Aerogravity Assist at Titan and Triton for Capture into Orbit About Saturn and Neptune

Previous work by our group has shown that an aerogravity assist maneuver at the moon Titan could be used to capture a spacecraft into a closed orbit about Saturn if a nominal atmospheric profile at Titan is assumed. The present study extends that work and examines the impact of atmospheric dispersions, variations in the final target orbit and low density aerodynamics on the aerocapture maneuver. Accounting for atmospheric dispersions substantially reduces the entry corridor width for a blunt configuration with a lift-to-drag ratio of 0.25. Moreover, the choice of the outbound hyperbolic excess speed (with respect to Titan) strongly influences the corridor width. Given the influence of these two parameters, certain mission scenarios may be feasible using a blunt aeroshell, while other mission designs would likely require a biconic vehicle with a higher lift-to-drag ratio. Preliminary simulations indicate that the same technique may be feasible for capture into orbit about Neptune using the tenuous atmosphere of Triton.

Ramsey, Philip↗

NASA NTRS2005

HyperPASS, a New Aeroassist Tool

A new software tool designed to perform aeroassist studies has been developed by Global Aerospace Corporation (GAC). The Hypersonic Planetary Aeroassist Simulation System (HyperPASS) [1] enables users to perform guided aerocapture, guided ballute aerocapture, aerobraking, orbit decay, or unguided entry simulations at any of six target bodies (Venus, Earth, Mars, Jupiter, Titan, or Neptune). HyperPASS is currently being used for trade studies to investigate (1) aerocapture performance with alternate aeroshell types, varying flight path angle and entry velocity, different gload and heating limits, and angle of attack and angle of bank variations; (2) variable, attached ballute geometry; (3) railgun launched projectile trajectories, and (4) preliminary orbit decay evolution. After completing a simulation, there are numerous visualization options in which data can be plotted, saved, or exported to various formats. Several analysis examples will be described.

Gates, Kristin↗

NASA NTRS2005

In-Space Propulsion (ISP) Aerocapture Technology

A viewgraph presentation is shown to raise awareness of aerocapture technology through in-space propulsion. The topics include: 1) Purpose; 2) In-Space Propulsion Program; 3) Aerocapture Overview; 4) Aerocapture Technology Alternatives; 5) Aerocapture Technology Project Process; 6) Results from 2002 Aerocapture TAG; 7) Bounding Case Requirements; 8) ST9 Flight Demonstration Opportunity; 9) Aerocapture NRA Content: Cycles 1 and 2; 10) Ames Research Center TPS Development; 11) Applied Research Associates TPS Development; 12) LaRC Structures Development; 13) Lockheed Martin Astronautics Aeroshell Development; 14) ELORET/ARC Sensor Development; 15) Ball Aerospace Trailing Ballute Development; 16) Cycle 2 NRA Selections - Aerocapture; and 17) Summary.

Munk, Michelle M.↗

NASA NTRS2007

Mars Science Laboratory: Entry, Descent, and Landing System Performance

In 2010, the Mars Science Laboratory (MSL) mission will pioneer the next generation of robotic Entry, Descent, and Landing (EDL) systems, by delivering the largest and most capable rover to date to the surface of Mars. To do so, MSL will fly a guided lifting entry at a lift-to-drag ratio in excess of that ever flown at Mars, deploy the largest parachute ever at Mars, and perform a novel Sky Crane maneuver. Through improved altitude capability, increased latitude coverage, and more accurate payload delivery, MSL is allowing the science community to consider the exploration of previously inaccessible regions of the planet. The MSL EDL system is a new EDL architecture based on Viking heritage technologies and designed to meet the challenges of landing increasing massive payloads on Mars. In accordance with level-1 requirements, the MSL EDL system is being designed to land an 850 kg rover to altitudes as high as 1 km above the Mars Orbiter Laser Altimeter defined areoid within 10 km of the desired landing site. Accordingly, MSL will enter the largest entry mass, fly the largest 70 degree sphere-cone aeroshell, generate the largest hypersonic lift-to-drag ratio, and deploy the largest Disk-Gap-Band supersonic parachute of any previous mission to Mars. Major EDL events include a hypersonic guided entry, supersonic parachute deploy and inflation, subsonic heatshield jettison, terminal descent sensor acquisition, powered descent initiation, sky crane terminal descent, rover touchdown detection, and descent stage flyaway. Key performance metrics, derived from level-1 requirements and tracked by the EDL design team to indicate performance capability and timeline margins, include altitude and range at parachute deploy, time on radar, and propellant use. The MSL EDL system, which will continue to develop over the next three years, will enable a notable extension in the advancement of Mars surface science by delivering more science capability than ever before to the surface of Mars. This paper describes the current MSL EDL system performance as predicted by end-to-end EDL simulations, highlights the sensitivity of this baseline performance to several key environmental assumptions, and discusses some of the challenges faced in delivering such an unprecedented rover payload to the surface of Mars.

Way, David W.↗

NASA NTRS2007

Inflatable Aerocapture Decelerators for Mars Orbiters

A multi-disciplinary research program was recently completed, sponsored by NASA Marshall Space Flight Center, on the subject of aerocapture of spacecraft weighing up to 5 metric tons at Mars. Heavier spacecraft will require deployable drag area beyond the dimensional limits of current and planned launch fairings. This research focuses on the approach of lightweight inflatable decelerators constructed with thin films, using fiber reinforcement and having a temperature limitation of 500 C. Trajectory analysis defines trajectories for a range of low ballistic coefficients for which convective heat flux is compatible with the material set. Fluid-Structure Interaction (FSI) tools are expanded to include the rarified flow regime. Several non-symmetrical configurations are evaluated for their capability to develop lift as part of the necessary trajectory control strategy. Manufacturing technology is developed for 3-D stretch forming of polyimide films and for tailored fiber reinforcement of thin films. Finally, the mass of the decelerator is estimated and compared to the mass of a traditional rigid aeroshell.

Brown, Glen J.↗

NASA NTRS2006

Can We Power Future Mars Missions?

The Vision for Space Exploration identified the exploration of Mars as one of the key pathways. In response, NASAs Mars Program Office is developing a detailed mission lineup for the next decade that would lead to future explorations. Mission architectures for the next decade include both orbiters and landers. Existing power technologies, which could include solar panels, batteries, radioisotope power systems, and in the future fission power, could support these missions. Second and third decade explorations could target human precursor and human in-situ missions, building on increasingly complex architectures. Some of these could use potential feed forward from earlier Constellation missions to the Moon, discussed in the ESAS study. From a potential Mars Sample Return mission to human missions the complexity of the architectures increases, and with it the delivered mass and power requirements also amplify. The delivered mass at Mars mostly depends on the launch vehicle, while the landed mass might be further limited by EDL technologies, including the aeroshell, parachutes, landing platform, and pinpoint landing. The resulting in-situ mass could be further divided into payload elements and suitable supporting power systems. These power systems can range from tens of watts to multi-kilowatts, influenced by mission type, mission configuration, landing location, mission duration, and season. Regardless, the power system design should match the power needs of these surface assets within a given architecture. Consequently, in this paper we will identify potential needs and bounds of delivered mass and architecture dependent power requirements to surface assets that would enable future in-situ exploration of Mars.

solar panels↗

NASA NTRS2008

NASA's In-Space Propulsion Technology Project Overview, Near-term Products and Mission Applicability

The In-Space Propulsion Technology (ISPT) Project, funded by NASA's Science Mission Directorate (SMD), is continuing to invest in propulsion technologies that will enable or enhance NASA robotic science missions. This overview provides development status, near-term mission benefits, applicability, and availability of in-space propulsion technologies in the areas of aerocapture, electric propulsion, advanced chemical thrusters, and systems analysis tools. Aerocapture investments improved (1) guidance, navigation, and control models of blunt-body rigid aeroshells, 2) atmospheric models for Earth, Titan, Mars and Venus, and 3) models for aerothermal effects. Investments in electric propulsion technologies focused on completing NASA s Evolutionary Xenon Thruster (NEXT) ion propulsion system, a 0.6-7 kW throttle-able gridded ion system. The project is also concluding its High Voltage Hall Accelerator (HiVHAC) mid-term product specifically designed for a low-cost electric propulsion option. The primary chemical propulsion investment is on the high-temperature Advanced Material Bipropellant Rocket (AMBR) engine providing higher performance for lower cost. The project is also delivering products to assist technology infusion and quantify mission applicability and benefits through mission analysis and tools. In-space propulsion technologies are applicable, and potentially enabling for flagship destinations currently under evaluation, as well as having broad applicability to future Discovery and New Frontiers mission solicitations.

Dankanich, John↗

NASA NTRS2006

Mars Science Laboratory: Entry, Descent, and Landing System Performance

Way, David W.↗

NASA NTRS2009

Ultra-high Temperature Emittance Measurements for Space and Missile Applications

Advanced modeling and design efforts for many aerospace components require high temperature emittance data. Applications requiring emittance data include propulsion systems, radiators, aeroshells, heatshields/thermal protection systems, and leading edge surfaces. The objective of this work is to provide emittance data at ultra-high temperatures. MSFC has a new instrument for the measurement of emittance at ultra-high temperatures, the Ultra-High Temperature Emissometer System (Ultra-HITEMS). AZ Technology Inc. developed the instrument, designed to provide emittance measurements over the temperature range 700-3500K. The Ultra-HITEMS instrument measures the emittance of samples, heated by lasers, in vacuum, using a blackbody source and a Fourier Transform Spectrometer. Detectors in a Nicolet 6700 FT-IR spectrometer measure emittance over the spectral range of 0.4-25 microns. Emitted energy from the specimen and output from a Mikron M390S blackbody source at the same temperature with matched collection geometry are measured. Integrating emittance over the spectral range yields the total emittance. The ratio provides a direct measure of total hemispherical emittance. Samples are heated using lasers. Optical pyrometry provides temperature data. Optical filters prevent interference from the heating lasers. Data for Inconel 718 show excellent agreement with results from literature and ASTM 835. Measurements taken from levitated spherical specimens provide total hemispherical emittance data; measurements taken from flat specimens mounted in the chamber provide near-normal emittance data. Data from selected characterization studies will be presented. The Ultra-HITEMS technique could advance space and missile technologies by advancing the knowledge base and the technology readiness level for ultra-high temperature materials.

Rogers, Jan↗

NASA NTRS2009

The NASA In-Space Propulsion Technology Project, Products, and Mission Applicability

The In-Space Propulsion Technology (ISPT) Project, funded by NASA s Science Mission Directorate (SMD), is continuing to invest in propulsion technologies that will enable or enhance NASA robotic science missions. This overview provides development status, near-term mission benefits, applicability, and availability of in-space propulsion technologies in the areas of aerocapture, electric propulsion, advanced chemical thrusters, and systems analysis tools. Aerocapture investments improved: guidance, navigation, and control models of blunt-body rigid aeroshells; atmospheric models for Earth, Titan, Mars, and Venus; and models for aerothermal effects. Investments in electric propulsion technologies focused on completing NASA s Evolutionary Xenon Thruster (NEXT) ion propulsion system, a 0.6 to 7 kW throttle-able gridded ion system. The project is also concluding its High Voltage Hall Accelerator (HiVHAC) mid-term product specifically designed for a low-cost electric propulsion option. The primary chemical propulsion investment is on the high-temperature Advanced Material Bipropellant Rocket (AMBR) engine providing higher performance for lower cost. The project is also delivering products to assist technology infusion and quantify mission applicability and benefits through mission analysis and tools. In-space propulsion technologies are applicable, and potentially enabling for flagship destinations currently under evaluation, as well as having broad applicability to future Discovery and New Frontiers mission solicitations.

Anderson, David J.↗

NASA NTRS2009

Aerocapture Guidance and Performance at Mars for High-Mass Systems

The Mars Entry, Descent, and Landing System Analysis (EDL-SA) project has been tasked with performing systems analysis to identify the optimal technologies required to land a 20-50 MT Exploration-class mission on Mars. It has been shown that it is not possible to safely land these large systems using heritage Mars EDL systems, or analogous Earth or Moon EDL systems. In 2007, NASA conducted a Mars Exploration Architecture Study which included an in depth review of the science motivations and engineering technology requirements for a human Mars mission campaign. This study resulted in an update to the Mars Design Reference Architecture (DRA 5.0). Among the primary findings and recommendations was the conclusion that landing of large payloads (greater than 1 MT) on the surface of Mars remains a key architectural challenge. Additionally, research and system studies of fundamental EDL technologies were highly recommended. The EDL-SA project identified the candidate technologies and assembled them into full capture and EDL sequences so that simulations could be developed to evaluate them. The chosen architectures combine the various technologies of interest in eight different ways. For aerocapture, only two scenarios were considered. The first is a rigid mid-L/D aeroshell (AS), which is represented in architectures 1, 4, 5, and 7. This scenario calls for a vehicle that flies at a 55-degree angle of attack, resulting in ballistic coefficient and L/D values of 490 kg/m 2 and 0.43, respectively. The second is a lifting hypersonic inflatable aerodynamic decelerator (HIAD), which is represented in architectures 2, 6, and 8. This scenario requires the vehicle to fly at a 22.2- degree angle of attack, which correlates to an L/D of 0.3 and was sized to provide a ballistic coefficient of 165 kg/m 2 . Architecture 3 is an all-propulsive sequence, and is not considered in this study.

Zumwalt, Carlie H.↗