SEARCH · Search NASA
Search indexed NASA NTRS and DOE OSTI research on propulsion, heat transfer, battery materials and energy systems. Follow report and document links to the original sources.
Quote a phrase for an exact phrase match. Source license links do not imply unrestricted reuse.
Report describes Simulator for Imager for Mars Pathfinder (SIMP) computer program. SIMP generates "virtual reality" display of view through video camera on Mars lander spacecraft of Mars Pathfinder mission, along with display of pertinent textual and graphical data, for use by scientific investigators in planning sequences of activities for mission.
Exchanging heat between hot and cold streams within an advanced life support system can save energy. This savings will reduce the equivalent system mass (ESM) of the system. Different system configurations are examined under steady-state conditions for various percentages of food growth and waste treatment. The scenarios investigated represent possible design options for a Mars reference mission. Reference mission definitions are drawn from the ALSS Modeling and Analysis Reference Missions Document, which includes definitions for space station evolution, Mars landers, and a Mars base. For each scenario, streams requiring heating or cooling are identified and characterized by mass flow, supply and target temperatures and heat capacities. The Pinch Technique is applied to identify good matches for energy exchange between the hot and cold streams and to calculate the minimum external heating and cooling requirements for the system. For each pair of hot and cold streams that are matched, there will be a reduction in the amount of external heating and cooling required, and the original heating and cooling equipment will be replaced with a heat exchanger. The net cost savings can be either positive or negative for each stream pairing, and the priority for implementing each pairing can be ranked according to its potential cost savings. Using the Pinch technique, a complete system heat exchange network is developed and heat exchangers are sized to allow for calculation of ESM. The energy-integrated design typically has a lower total ESM than the original design with no energy integration. A comparison of ESM savings in each of the scenarios is made to direct future Pinch Analysis efforts.
The Viking Lander 1 and 2 cameras acquired many high-resolution pictures of the Chryse Planitia and Utopia Planitia landing sites. Based on computer-processed data of a selected number of these pictures, eight high-resolution mosaics were published by the U.S. Geological Survey as part of the Atlas of Mars, Miscellaneous Investigation Series. The mosaics are composites of the best picture elements (pixels) of all the Lander pictures used. Each complete mosaic extends 342.5 deg in azimuth, from approximately 5 deg above the horizon to 60 deg below, and incorporates approximately 15 million pixels. Each mosaic is shown in a set of five sheets. One sheet contains the full panorama from one camera taken in either morning or evening. The other four sheets show sectors of the panorama at an enlarged scale; when joined together they make a panorama approximately 2' X 9'.
Phenolic Impregnated Carbon Ablator (PICA) is a low-density ablator that has been used as the planetary entry heatshield for several NASA missions since the late 90’s. Its low density and efficient performance characteristics have proven effective for use from Discovery to Flag-ship class missions, from Sample Return missions such as Stardust, OSIRIS-REx to large Mars Lander missions such as Mars Science Lab (MSL) and Mars 2020. The rayon-based carbon precursor raw material used in PICA preform manufacturing has experienced multiple supply chain issues. The challenge involved in finding a replacement fiber source is in processing as well as in the final performance of the ablator. Each replacement necessitates the requalification of the PICA. This has happened at least twice in the past 25 years, and a third substitution is now needed. Due to the obsolescence of the input foreign rayon fiber source, a new variant of PICA has been developed using a domestic rayon-like fiber source, Lyocell. Due to its flight heritage and proven capability, PICA is baselined as the Thermal Protection System (TPS) for Dragonfly and Mars Sample Return Sample Retrieval Lander mission and is being considered for the backshell of the Mars Sample Return Earth Entry System. All three missions are due to be launched between 2026 and 2028. The challenge this time is to ensure the PICA made with domestic material is a suitable replacement to the heritage PICA used in MSL and Mars 2020 so that the design of the heatshield can be matured without much risk. Results are presented from the recent efforts of 22 PICA-D billets that were Lot Acceptance Tested. Thermal, mechanical, and representative environment arc-jet tests have been conducted. Testing of PICA-Domestic (PICA-D) indicates very comparable performance with respect to “heritage” PICA materials and thus PICA-D is expected to be a sustainable and nearly a “drop-in” replacement solution for future NASA missions.
Phenolic Impregnated Carbon Ablator (PICA) is a low-density ablator that has been used as the planetary entry heatshield for several NASA missions since the late 90’s. Its low density and efficient performance characteristics have proven effective for use from Discovery to Flag-ship class missions, from Sample Return missions such as Stardust, OSIRIS-REx to large Mars Lander missions such as Mars Science Lab (MSL) and Mars 2020. The rayon-based carbon precursor raw material used in PICA preform manufacturing has experienced multiple supply chain issues. The challenge involved in finding a replacement fiber source is in processing as well as in the final performance of the ablator. Each replacement necessitates the requalification of the PICA. This has happened at least twice in the past 25 years, and a third substitution is now needed. Due to the obsolescence of the input foreign rayon fiber source, a new variant of PICA has been developed using a domestic rayon-like fiber source, Lyocell. Due to its flight heritage and proven capability, PICA is baselined as the Thermal Protection System (TPS) for Dragonfly and Mars Sample Return Sample Retrieval Lander mission and is being considered for the backshell of the Mars Sample Return Earth Entry System. All three missions are due to be launched between 2026 and 2028. The challenge this time is to ensure the PICA made with domestic material is a suitable replacement to the heritage PICA used in MSL and Mars 2020 so that the design of the heatshield can be matured without much risk. Results are presented from the recent efforts of 22 PICA-D billets that were Lot Acceptance Tested. Thermal, mechanical, and representative environment arc-jet tests have been conducted. Testing of PICA-Domestic (PICA-D) indicates very comparable performance with respect to “heritage” PICA materials and thus PICA-D is expected to be a sustainable and nearly a “drop-in” replacement solution for future NASA missions.
Supporting data for mission planning and design criteria of Mars hard landing system
NASA’s successful Mars Atmosphere and Volatile Evolution Mission (MAVEN), currently engaged in a survey of the Martian atmosphere and its evolution over time, has successfully completed a series of orbit changes, including a two-month aerobraking campaign, to satisfy additional mission roles. Launched on 18 November 2013, it entered Martian orbit on 22 September 2014. Having finished its primary one-year mission, the NASA Mars Program Office (MPO) started discussions with the MAVEN project on feasible options for extending the mission lifetime along with providing support for various Mars lander functions. MAVEN is the latest of the three active NASA Mars orbiters, with no others planned for the foreseeable future. Given that the landers need orbital relay support for years to come, and the MRO and Odyssey orbiters are aging, MAVEN’s role in providing critical orbital relay support has become increasingly important. After several years of studies, MPO directed the MAVEN project in 2018 to reduce its orbit size to provide better telecommunications relay support for Martian surface landers and to support Mars 2020 EDL under the constraint that MAVEN will have enough fuel to operate until 2030. As a part of this effort, the recent aerobraking campaign was one of several steps crucial to accomplishing this goal. This paper discusses the challenges of the entire mission design and its distillation into a concrete set of targeted objectives. With the successful orbit change, MAVEN will be able to provide support for the Mars 2020 relay activities from EDL through surface operations beyond 2030. This effort can serve as an excellent example for future missions having similar needs and requirements.
Accumulation of wind-blown dust particles on solar cells and instruments will be a great challenge in the exploration of Mars, significantly reducing their lifetime, durability, and power output. For future Mars Lander missions it is crucial to gain information about the ideal angle at which solar panels can be positioned to minimize dust deposition and thus, maximize the power output and lifetime of the solar cells. The major determinant for the optimal panel angle is the angle of repose of the dust particles that is dependent on a variety of physical and chemical properties of the particles, the panel surface, and the environmental conditions on the Mars surface. To gain a basic understanding of the physical and chemical processes that govern dust deposition and to get feedback for the design of an experiment suitable for one of the future Mars Lander missions we simulate atmospheric conditions expected on the Mars surface in a controlled chamber, and observe the angle of repose of Mars dust surrogates. Dust deposition and angle of repose were observed on different sized spheres. To cover a range of potential materials we will use spheres made of 7075 aluminum (10 mm, and 15 mm), alumina oxide ceramic (10 mm), and Teflon(trademark) (10 mm) and wafers of gallium arsenide, silicon.
Under existing COSPAR policy adopted in 1984, missions to Mars (landers, probes, and some orbiters) are designated as Category IV missions. As such, the procedures for implementing planetary protection requirements could include trajectory biasing, cleanrooms, bioload reduction, sterilization of hardware, and bioshields. In 1992, a U.S. National Research Council study recommended that controls on forward contamination of Mars be tied to specific mission objectives. The report recommended that Mars landers with life detection instruments be subject to at least Viking-level sterilization procedures for bioload reduction, while spacecraft (including orbiters) without life detection instruments be subject to at least Viking-level pre-sterilization procedures for bioload reduction but need not be sterilized. In light of this, it is proposed that the current policy's Category IV and its planetary protection requirements be divided into two sub-categories as follows: Category IVa, for missions comprising landers and probes without life detection experiments, which will meet a specified bioburden limit for exposed surfaces, and Category IVb, for landers and probes with life detection experiments, which will require sterilization of landed systems. In addition, Category III orbiter mission specifications are expanded to be consistent with these recommendations.
NASA s Phoenix Mars Lander began its journey to Mars from Cape Canaveral, Florida in August 2007, but its journey to the launch pad began many years earlier in 1997 as NASA s Mars Surveyor Program 2001 Lander. In the intervening years, the entry, descent and landing (EDL) system architecture went through a series of changes, resulting in the system flown to the surface of Mars on May 25th, 2008. Some changes, such as entry velocity and landing site elevation, were the result of differences in mission design. Other changes, including the removal of hypersonic guidance, the reformulation of the parachute deployment algorithm, and the addition of the backshell avoidance maneuver, were driven by constant efforts to augment system robustness. An overview of the Phoenix EDL system architecture is presented along with rationales driving these architectural changes.
Unmanned planetary probe and lander capsule design, emphasizing Mars lander capsules
Under existing COSPAR policy adopted in 1984, missions to Mars (landers, probes, and some orbiters) are designated as Category IV missions. As such, the procedures for implementing planetary protection requirements could include trajectory biasing, cleanrooms, bioload reduction, sterilization of hardware, and bioshields, i. e. requirements could be similar to Viking. However, in 1992, a U. S. National Academy of Sciences study recommended that controls on forward contamination of Mars be tied to specific mission objectives. The report recommended that Mars landers with life detection instruments be subject to at least Viking-level sterilization procedures for bioload reduction, while spacecraft (including orbiters) without life detection instruments be subject to at least Viking-level pre sterilization procedures for bioload reduction but need not be sterilized. In light of this, it is proposed that the current policy's Category IV missions and their planetary protection requirements be divided into two subcategories as follows: Category IV A, for missions comprising landers and probes without life detection experiments and some orbiters, which will meet a specified bioburden limit for exposed surfaces; Category IV B, for landers and probes with life detection experiments, which will require complete system sterilization. For Category IV A missions, bioburden specifications will be proposed and implementing procedures discussed. A resolution will be proposed to modify the existing COSPAR policy to reflect these changes. Similar specifications, procedures, and resolution for Category IV B missions will be the subject of a later study.
The objectives of this program are to 1) Assess viability of using lithium-ion technology for future NASA applications, with emphasis upon Mars landers and rovers which will operate on the planetary surface; 2) Support the JPL 2003 Mars Exploration Rover program to assist in the delivery and testing of a 8 AHr Lithium-Ion battery (Lithion/Yardney) which will power the rover; 3) Demonstrate applicability of using lithium-ion technologyfor future Mars applications: Mars 09 Science Laboratory (Smart Lander) and Future Mars Surface Operations (General). Mission simulation testing was carried out for cells and batteries on the Mars Surveyor 2001 Lander and the 2003 Mars Exploration Rover.
Development of a powered descent capability for atmospheric environments is heavily reliant on computational simulation. The prohibitive computational cost of such simulations motivates an improvement in the understanding of the minimum computational fidelity re-quired to accurately characterize aerodynamic-propulsive interference for such applications. This work examines the applicability of detached eddy simulation methods for retropropulsion in atmospheric environments through utilization of a GPU-accelerated computational framework, yielding data that are largely unachievable with conventional high-performance computing resources. This effort was specifically designed to quantitatively assess the effects of spatial resolution on vehicle aerodynamics for nominal operation of a low lift-to-drag ratio, human-scale Mars lander concept. The test matrix and scaling approach span relevant nozzle expansion conditions as well as mid-supersonic to high-subsonic operating conditions. Solutions were generated using computational grids ranging from 143 million to 1.14 billion grid points (degrees of freedom). This paper will provide an overview of the computational campaign, approach, and discussion of preliminary results focused on a range of operating conditions for a conceptual low lift-to-drag, human-scale Mars lander.
The Mars Pathfinder (MPF) Spacecraft was built and tested at the Jet Propulsion Laboratory during 1995/96. MPF is scheduled to launch in December 1996 and to land on Mars on July 4, 1997. The testing program for MPF required subjecting the mission hardware to both deep space and Mars surface conditions. A series of tests were devised and conducted from 1/95 to 7/96 to study the thermal response of the MPF spacecraft to the environmental conditions in which it will be exposed during the cruise phase (on the way to Mars) and the lander phase (landed on Mars) of the mission. Also, several tests were conducted to study the thermal characteristics of the Mars rover, Sojourner, under Mars surface environmental conditions. For these tests, several special test fixtures and methods were devised to simulate the required environmental conditions. Creating simulated Mars surface conditions was a challenging undertaking since Mars' surface is subjected to diurnal cycling between -20 C and -85 C, with windspeeds to 20 m/sec, occurring in an 8 torr CO2 atmosphere. This paper describes the MPF test program which was conducted at JPL to verify the MPF thermal design.
Landing site dispersions of unmanned spacecraft on Mars
In fiscal year 1997, NASA, the Jet Propulsion Laboratory, and the U.S. Air Force established a joint program to competitively develop high-power, rechargeable lithium-ion battery technology for aerospace applications. The goal was to address Department of Defense and NASA requirements not met by commercial battery developments. Under this program, contracts have been awarded to Yardney Technical Products, Eagle- Picher Technologies, LLC, BlueStar Advanced Technology Corporation, and SAFT America, Inc., to develop cylindrical and prismatic cell and battery systems for a variety of NASA and U.S. Air Force applications. The battery systems being developed range from low-capacity (7 to 20 A-hr) and low-voltage (14 to 28 V) systems for planetary landers and rovers to systems for aircraft that require up to 270 V and for Unmanned Aerial Vehicles that require capacities up to 200 A-hr. Low-Earth-orbit and geosynchronousorbit spacecraft pose additional challenges to system operation with long cycle life (>30,000 cycles) and long calendar life (>10 years), respectively.