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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.

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NASA NTRSDate not supplied

A CFD Validation Study using Artemis 1 Orbital Slosh Test Data

Advancements in liquid propellant management science and technologies are key to increasing safety, decreasing cost, and increasing payload mass of NASA space missions. Liquid propellant usually comprises a large portion of the total mass of launch vehicles and spacecraft, so predicting and controlling the motion of it are important. Computational fluid dynamics (CFD) programs are critical to predicting slosh dynamics, but CFD programs require experimental validation with physically relevant test data before the results can be trusted. Low-gravity slosh test data is lacking, and most of what exists is inadequate for CFD validation. A potential slosh risk for Orion, and the lack of validated low-gravity slosh models, led to performing on-orbit slosh tests during the Artemis 1 mission to assess the impacts of slosh on the guidance, navigation, and control of Orion. This work details the validation of a coupled slosh-motion CFD model using the Orion orbital slosh test data set.

CFD↗

NASA NTRSDate not supplied

Structural Analysis of Load-Bearing Components in Mars Science Helicopter

The successful demonstration of powered flight on Mars by Ingenuity has led to the development of next generation Martian rotorcraft. As technology advances from a technology demonstrator to possible expanded planetary science investigations, the potential to fly a high payload carrying rotorcraft relies on the capability to analyze and validate the structural integrity of the vehicle in addition to the rotor aerodynamic performance. The Mars Science Helicopter (MSH), a next generation hexacopter concept, is one example of a proposed vehicle that would perform science independently of a land-based vehicle. To ensure the feasibility and viability of MSH mission performance, it is critical to mature the structural design for these vehicle concepts to bridge the gap between the best practices of the spacecraft and aircraft communities. MSH was analyzed under operational and 26G quasi-static launch loads with different rotor arm and fuselage designs. Structural analysis results showed that rotor arms could be stiffened for significantly improved structural performance with minimal mass penalty, while the rotorcraft frame exhibited a more complex relationship between mass and launch load stresses due to the mass-dependent launch loads.

Load-Bearing↗

NASA NTRS2025

Structural Analysis of Load-Bearing Components in Mars Science Helicopter

The successful demonstration of powered flight on Mars by Ingenuity has led to the development of next generation Martian rotorcraft. As technology advances from a technology demonstrator to possible expanded planetary science investigations, the potential to fly a high payload carrying rotorcraft relies on the capability to analyze and validate the structural integrity of the vehicle in addition to the rotor aerodynamic performance. The Mars Science Helicopter (MSH), a next generation hexacopter concept, is an example of a proposed vehicle that would perform science investigations independent of a land-based vehicle. To ensure the feasibility and viability of MSH mission performance, it is critical to mature the structural design of vehicle concepts to bridge the gap between the best practices of the spacecraft and aircraft communities. MSH was analyzed under operational and 26G quasi-static launch loads with different rotor arm and frame designs. Structural analysis results showed that rotor arms could be stiffened for significantly improved structural performance with minimal mass penalty, while the rotorcraft frame exhibited a more complex relationship between mass and launch load stresses due to the mass-dependent launch loads.

Vehicle↗

NASA NTRSDate not supplied

A CFD Validation Study Using Artemis 1 Orbital Slosh Test Data

CFD↗

NASA NTRS2001

Optical Design of the MOSES Sounding Rocket Experiment

The Multi-Order Solar EUV Spectrograph (MOSES) is a sounding rocket payload now being developed by Montana State University in collaboration with the Goddard Space Flight Center, Lockheed Martin Advanced Technology Center, and Mullard Space Science Laboratory. The instrument utilizes a unique optical design to provide solar EUV measurements with true 2-pixel resolutions of 1.0 arcsec and 60 mA over a full two-dimensional field of view of 1056 x 528 arcsec, all at a time cadence of 10 s. This unprecedented capability is achieved by means of an objective spherical grating 100 mm in diameter, ruled at 833 gr/mm. The concave grating focuses spectrally dispersed solar radiation onto three separate detectors, simultaneously recording the zero-order as well as the plus and minus first-spectral-order images. Data analysis procedures, similar to those used in X-ray tomography reconstructions, can then disentangle the mixed spatial and spectral information recorded by the multiple detectors. A flat folding mirror permits an imaging focal length of 4.74 m to be packaged within the payload's physical length of 2.82 m. Both the objective grating and folding flat have specialized, closely matched, multilayer coatings that strongly enhance their EUV reflectance while also suppressing off-band radiation that would otherwise complicate data inversion. Although the spectral bandpass is rather narrow, several candidate wavelength intervals are available to carry out truly unique scientific studies of the outer solar atmosphere. Initial flights of MOSES, scheduled to begin in 2004, will observe a 10 Angstrom band that covers very strong emission lines characteristic of both the sun's corona (Si XI 303 Angstroms) and transition-region (He II 304 Angstroms). The MOSES program is supported by a grant from NASA's Office of Space Science.

Thomas, Roger J.↗

NASA NTRS2007

Solar Electric and Chemical Propulsion Technology Applications to a Titan Orbiter/Lander Mission

Several advanced propulsion technology options were assessed for a conceptual Titan Orbiter/Lander mission. For convenience of presentation, the mission was broken into two phases: interplanetary and Titan capture. The interplanetary phase of the mission was evaluated for an advanced Solar Electric Propulsion System (SEPS), while the Titan capture phase was evaluated for state-of-art chemical propulsion (NTO/Hydrazine), three advanced chemical propulsion options (LOX/Hydrazine, Fluorine/Hydrazine, high Isp mono-propellant), and advanced tank technologies. Hence, this study was referred to as a SEPS/Chemical based option. The SEPS/Chemical study results were briefly compared to a 2002 NASA study that included two general propulsion options for the same conceptual mission: an all propulsive based mission and a SEPS/Aerocapture based mission. The SEP/Chemical study assumed identical science payload as the 2002 NASA study science payload. The SEPS/Chemical study results indicated that the Titan mission was feasible for a medium launch vehicle, an interplanetary transfer time of approximately 8 years, an advanced SEPS (30 kW), and current chemical engine technology (yet with advanced tanks) for the Titan capture. The 2002 NASA study showed the feasibility of the mission based on a somewhat smaller medium launch vehicle, an interplanetary transfer time of approximately 5.9 years, an advanced SEPS (24 kW), and advanced Aerocapture based propulsion technology for the Titan capture. Further comparisons and study results were presented for the advanced chemical and advanced tank technologies.

Cupples, Michael↗

NASA NTRS1972

Study of the application of advanced technologies to long-range transport aircraft. Volume 2: Research and development requirements

Investigations were conducted to evaluate the relative benefits attainable through the exploitation of advanced technologies and to identify future research and development efforts required to permit the application of selected technologies to transport aircraft entering commercial operation in 1985. Results show that technology advances, particularly in the areas of composite materials, supercritical aerodynamics, and active control systems, will permit the development of long-range, high-payload commercial transports operating at high-subsonic speeds with direct operating costs lower than those of current aircraft. These advanced transports also achieve lower noise levels and lower engine pollutant emissions than current transports. Research and development efforts, including analytical investigations, laboratory test programs, and flight test programs, are required in essentially all technology areas to achieve the potential technology benefits.

Lange, R. H.↗

NASA NTRSDate not supplied

Mars Science Helicopter: Conceptual Design of the Next Generation of Mars Rotorcraft

Robotic planetary aerial vehicles increase the range of terrain that can be examined, compared to traditional landers and rovers, and have more near-surface capability than orbiters. Aerial mobility is a promising possibility for planetary exploration as it reduces the challenges that difficult obstacles pose to ground vehicles. The first use of a rotorcraft for a planetary mission will be in 2021, when the Ingenuity Mars helicopter technology demonstrator will be deployed via the Perseverance rover. NASA’s Jet Propulsion Laboratory and NASA Ames Research Center are exploring possibilities for a Mars Science Helicopter, a second-generation Mars rotorcraft with the capability of conducting science investigations independently of a lander or rover (although this type of vehicle could also be used to assist rovers or landers in future missions). Two, large rotorcraft configurations are described: a hexacopter and a co-axial helicopter with a payload in the range of two to three kilograms and an overall vehicle mass of approximately twenty kilograms. Additionally, advancements in technology over the course of the study are applied to a rotorcraft of the same size and form as Ingenuity. Initial estimates of weight and performance were based on the capabilities of Ingenuity. Rotorcraft designs for Mars are constrained by the dimensions of the aeroshell and lander for the trip to the planet, constraining maximum rotor dimensions and, hence, overall performance potential. The effects of airfoils designed specifically for the low Reynolds number and high Mach number inherent to operation on Mars were studied. Rotor structural designs were developed that met blade frequency and weight targets, subject to material stress limits. The final designs are representative of the vehicle configurations required for a large range of future missions and will require relatively minor adaptations once science tasks are chosen. These designs will be compared to Ingenuity to demonstrate technology advancements developed during the study.

Mars Science Helicopter↗

NASA NTRS1989

Laser-boosted lightcraft technology demonstrator

The ultimate goal for this NASA/USRA-sponsored 'Apollo Lightcraft Project' is to develop a revolutionary manned launch vehicle technology that can potentially reduce payload transport costs by a factor of 1000 below the space shuttle orbiter. The Rensellaer design team proposes to utilize advanced, highly energetic, beamed-energy sources (laser, microwave) and innovative combined-cycle (airbreathing/rocket) engines to accomplish this goal. This year's effort, the detailed description and performance analysis of an unmanned 1.4-m Lightcraft Technology Demonstrator (LTD) drone, is presented. The novel launch system employs a 100-MW-class ground-based laser to transmit power directly to an advanced combined-cycle engine that propels the 120-kg LTD to orbit, with a mass ratio of two. The single-stage-to-orbit (SSTO) LTD machine then becomes an autonomous sensor satellite that can deliver precise, high-quality information typical of today's large orbital platforms. The dominant motivation behind this study is to provide an example of how laser propulsion and its low launch costs can induce a comparable order-of-magnitude reduction in sensor satellite packaging costs. The issue is simply one of production technology for future, survivable SSTO aerospace vehicles that intimately share both laser propulsion engine and satellite functional hardware. A mass production cost goal of 10(exp 3)/kg for the LTD vehicle is probably realizable.

Antonison, M.↗

NASA NTRS1990

ATDRS payload technology R & D

Four technology development tasks were chosen to reduce (or at least better understand) the technology risks associated with proposed approaches to Advanced Tracking and Data Relay Satellite (ATDRS). The four tasks relate to a Tri-Band Antenna feed system, a Digital Beamforming System for the S Band Multiple-Access System (SMA), an SMA Phased Array Antenna, and a Configuration Thermal/Mechanical Analysis task. The objective, approach, and status of each are discussed.

Anzic, G.↗

NASA NTRS1989

Apollo Lightcraft Project

The ultimate goal for this NASA/USRA-sponsored Apollo Lightcraft Project is to develop a revolutionary manned launch vehicle technology which can potentially reduce payload transport costs by a factor of 1000 below the Space Shuttle Orbiter. The Rensselaer design team proposes to utilize advanced, highly energetic, beamed-energy sources (laser, microwave) and innovative combined-cycle (airbreathing/rocket) engines to accomplish this goal. The research effort focuses on the concept of a 100 MW-class, laser-boosted Lightcraft Technology Demonstrator (LTD) drone. The preliminary conceptual design of this 1.4 meter diameter microspacecraft involved an analytical performance analysis of the transatmospheric engine in its two modes of operation (including an assessment of propellant and tankage requirements), and a detailed design of internal structure and external aeroshell configuration. The central theme of this advanced propulsion research was to pick a known excellent working fluid (i.e., air or LN sub 2), and then to design a combined-cycle engine concept around it. Also, a structural vibration analysis was performed on the annular shroud pulsejet engine. Finally, the sensor satellite mission was examined to identify the requisite subsystem hardware: e.g., electrical power supply, optics and sensors, communications and attitude control systems.

Myrabo, Leik N.↗

NASA NTRS1972

Optimization of engines for a commercial Mach 0.98 transport using advanced turbine cooling methods

A study was made of an advanced technology airplane using supercritical aerodynamics. Cruise Mach number was 0.98 at 40,000 feet altitude with a payload of 60,000 pounds and a range of 3000 nautical miles. Separate-flow turbofans were examined parametrically to determine the effect of sea-level-static design turbine-inlet-temperature and noise on takeoff gross weight (TOGW) assuming full-film turbine cooling. The optimum turbine inlet temperature was 2650 F. Two-stage-fan engines, with cruise fan pressure ratio of 2.25, achieved a noise goal of 103.5 EPNdB with todays noise technology while one-stage-fan engines, achieved a noise goal of 98 EPNdB. The take-off gross weight penalty to use the one-stage fan was 6.2 percent.

Kraft, G. A.↗

NASA NTRS2013

High Altitude Supersonic Decelerator Test Vehicle

The Low Density Supersonic Decelerator (LDSD) project is tasked by NASA's Office of the Chief Technologist (OCT) to advance the state of the art in Mars entry and descent technology in order to allow for larger payloads to be delivered to Mars at higher altitudes with better accuracy. The project will develop a 33.5 m Do Supersonic Ringsail (SSRS) parachute, 6m attached torus, robotic class Supersonic Inflatable Aerodynamic Decelerator (SIAD-R), and an 8 m attached isotensoid, exploration class Supersonic Inflatable Aerodynamic Decelerator (SIAD-E). The SSRS and SIAD-R should be brought to TRL-6, while the SIAD-E should be brought to TRL-5. As part of the qualification and development program, LDSD must perform a Mach-scaled Supersonic Flight Dynamics Test (SFDT) in order to demonstrate successful free flight dynamic deployments at Mars equivalent altitude, of all three technologies. In order to perform these tests, LDSD must design and build a test vehicle to deliver all technologies to approximately 180,000 ft and Mach 4, deploy a SIAD, free fly to approximately Mach 2, deploy the SSRS, record high-speed and high-resolution imagery of both deployments, as well as record data from an instrumentation suite capable of characterizing the technology induced vehicle dynamics. The vehicle must also be recoverable after splashdown into the ocean under a nominal flight, while guaranteeing forensic data protection in an off nominal catastrophic failure of a test article that could result in a terminal velocity, tumbling water impact.

Cook, Brant T.↗

NASA NTRS1990

ATDRS payload technology research and development

Four technology development tasks were chosen to reduce (or at least better understand) the technology risks associated with proposed approaches to Advanced Tracking and Data Relay Satellite (ATDRS). The four tasks relate to a Tri-Band Antenna feed system, a Digital Beamforming System for the S Band Multiple Access System (SMA), an SMA Phased Array Antenna, and a Configuration Thermal/Mechanical Analysis task. The objective, approach, and status of each are discussed.

Anzic, G.↗

NASA NTRS2012

Composite Payload Fairing Structural Architecture Assessment and Selection

This paper provides a summary of the structural architecture assessments conducted and a recommendation for an affordable high performance composite structural concept to use on the next generation heavy-lift launch vehicle, the Space Launch System (SLS). The Structural Concepts Element of the Advanced Composites Technology (ACT) project and its follow on the Lightweight Spacecraft Structures and Materials (LSSM) project was tasked with evaluating a number of composite construction technologies for specific Ares V components: the Payload Shroud, the Interstage, and the Core Stage Intertank. Team studies strived to address the structural challenges, risks and needs for each of these vehicle components. Leveraging off of this work, the subsequent Composites for Exploration (CoEx) effort is focused on providing a composite structural concept to support the Payload Fairing for SLS. This paper documents the evaluation and down selection of composite construction technologies and evolution to the SLS Payload Fairing. Development of the evaluation criteria (also referred to as Figures of Merit or FOMs), their relative importance, and association to vehicle requirements are presented. A summary of the evaluation results, and a recommendation of the composite concept to baseline in the Composites for Exploration (CoEx) project is presented. The recommendation for the SLS Fairing is a Honeycomb Sandwich architecture based primarily on affordability and performance with two promising alternatives, Hat stiffened and Fiber Reinforced Foam (FRF) identified for eventual program block upgrade.

Krivanek, Thomas M.↗

NASA NTRS1972

A common solar electric upper stage for multi-mission applications.

The concept of using a solar-electric upper stage for multi-mission application is practical, economical, and versatile approach to space expMission capabilities include a Mercury orbiter, asteroid and comet rendezvous, a solar probe, and extra-ecliptic probe, and outer planet flybys and orbiters. The 17.5-kw vehicle has an injected mass of 1500 to 2500 kg and carries up to 500 kg of attached or separable payload packages. Mission characteristics, performance tradeoffs and a concepegn configuration are presented, and program plans, cost estimates and advanced technology requirements are identified.

Meissinger, H. F.↗

NASA NTRS1973

The space shuttle payload planning working groups. Volume 10: Space technology

The findings and recommendations of the Space Technology group of the space shuttle payload planning activity are presented. The elements of the space technology program are: (1) long duration exposure facility, (2) advanced technology laboratory, (3) physics and chemistry laboratory, (4) contamination experiments, and (5) laser information/data transmission technology. The space technology mission model is presented in tabular form. The proposed experiments to be conducted by each test facility are described. Recommended approaches for user community interfacing are included.

Source record↗

NASA NTRS1980

Air cargo market outlook and impact via the NASA CLASS project

An overview is given of the Cargo/Logistics Airlift Systems Study (CLASS) project which was a 10 man-year effort carried out by two contractor teams, aimed at defining factors impacting future system growth and obtaining market requirements and design guidelines for future air freighters. Growth projection was estimated by two approaches: one, an optimal systems approach with a more efficient and cost effective system considered as being available in 1990; and the other, an evolutionary approach with an econometric behavior model used to predict long term evolution from the present system. Both approaches predict significant growth in demand for international air freighter services and less growth for U.S. domestic services. Economic analysis of air freighter fleet options indicate very strong market appeal of derivative widebody transports in 1990 with little incentive to develop all new dedicated air freighters utilizing the 1990's technology until sometime beyond the year 2000. Advanced air freighters would be economically attractive for a wide range of payload sizes (to 500 metric tons), however, if a government would share in the RD and T costs by virtue of its needs for a slightly modified version of a civil air freighter design (a.g. military airlifter).

Winston, M. M.↗