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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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At least 253 records · Page 14
All sourcesNASA NTRSDOE OSTI
NASA NTRS1987

Mars manned transportation vehicle

A viable power system technology for a surface transportation vehicle to explore the planet Mars is presented. A number of power traction systems were investigated, and it was found that a regenerative hydrogen-oxygen fuel cell appears to be attractive for a manned Mars rover application. Mission requirements were obtained from the Manned Mars Mission Working Group. Power systems weights, power, and reactants requirements were determined as a function of vehicle weights for vehicles weighing from 6,000 to 16,000 lb (2,722 to 7,257 kg), (Earth weight). The vehicle performance requirements were: velocity, 10 km/hr; range, 100 km; slope climbing capability, 30 deg uphill for 50 km; mission duration, 5 days; and crew, 5. Power requirements for the operation of scientific equipment and support system capabilities were also specified and included in this study. The concept developed here would also be applicable to a Lunar based vehicle for Lunar exploration. The reduced gravity on the Lunar surface, (over that on the Martian surface), would result in an increased range or capability over that of the Mars vehicle since many of the power and energy requirements for the vehicle are gravity dependent.

Perez-Davis, Marla E.↗
NASA NTRS1989

Mars manned transportation vehicle

A viable power system technology for a surface transportation vehicle to explore the planet Mars is presented. A number of power traction systems were investigated, and it was found that a regenerative hydrogen-oxygen fuel cell appears to be attractive for a manned Mars rover application. Mission requirements were obtained from the Manned Mars Mission Working Group. Power systems weights, power, and reactants requirements were determined as a function of vehicle weights for vehicles weighting from 6,000 to 16,000 lb (2,722 to 7,257 kg), (Earth weight). The vehicle performance requirements were: velocity, 10 km/hr; range, 100 km; slope climbing capability, 30 deg uphill for 50 km; mission duration, 5 days; and crew, 5. Power requirements for the operation of scientific equipment and support system capabilities were also specified and included in this study. The concept developed here would also be applicable to a Lunar based vehicle for Lunar exploration. The reduced gravity on the Lunar surface, (over that on the Martian surface), would result in an increased range or capability over that of the Mars vehicle since many of the power and energy requirements for the vehicle are gravity dependent.

Perez-Davis, Marla E.↗
NASA NTRS2022

Treatment of Launch Vehicle Flight Control Stability Margin Reductions for Crewed Missions with Emphasis on Slosh Dynamics

NESC’s perspective for crewed spaceflight: Acceptance of flight control gain/phase stability margin reductions from industry standards should be accompanied by an adequately extensive technical treatment, including: •Analyzing the fundamental physics involved, with applicable simulation tool verification (particularly if results are dissimilar among rules of thumb, linear tools, nonlinear analysis, and flight data) •Conducting sensitivity studies in time and frequency domains to analyze effects of possible parameter and system variations •Studying the effects of the consequence of instability associated with offending modes by running stressing cases in time domain •Assessing alternative flight control designs to demonstrate that present design appropriately balances overall vehicle risk (i.e., quantitatively delineate chosen tradeoffs between various stability margins and vehicle performance in the context of risk/consequence) Work presented here represents an example summary of expected engineering work to flight-certify crewed missions with unstable slosh modes and reduced stability margins

Launch Vehicle Flight Control↗
NASA NTRS1981

Magnetic control systems for large spacecraft with applications to space telescope

Magnetic control systems for large space vehicles offer the advantage of a simple, reliable, low cost augmentation to the primary control system. When used for momentum management, a magnetic torque source offers a long life and noncontaminant environment when compared to a mass expulsion torque source. These qualities make such systems suitable for employment with the Space Telescope, which is a long life, high performance vehicle with optics and scientific instruments which would be degraded by contamination due to mass expulsion products. The various applications of magnetic systems on the Space Telescope are considered. The future trend in magnetic control of large space vehicles lies in providing a known three axis reference for backup operations, such as recovery of the primary control mode.

Dougherty, H.↗
NASA NTRS1982

Free-piston Stirling hydraulic engine and drive system for automobiles

The calculated fuel economy for an automotive free piston Stirling hydraulic engine and drive system using a pneumatic accumulator with the fuel economy of both a conventional 1980 spark ignition engine in an X body class vehicle and the estimated fuel economy of a 1984 spark ignition vehicle system are compared. The results show that the free piston Stirling hydraulic system with a two speed transmission has a combined fuel economy nearly twice that of the 1980 spark ignition engine - 21.5 versus 10.9 km/liter (50.7 versus 25.6 mpg) under comparable conditions. The fuel economy improvement over the 1984 spark ignition engine was 81 percent. The fuel economy sensitivity of the Stirling hydraulic system to system weight, number of transmission shifts, accumulator pressure ratio and maximum pressure, auxiliary power requirements, braking energy recovery, and varying vehicle performance requirements are considered. An important finding is that a multispeed transmission is not required. The penalty for a single speed versus a two speed transmission is about a 12 percent drop in combined fuel economy to 19.0 km/liter (44.7 mpg). This is still a 60 percent improvement in combined fuel economy over the projected 1984 spark ignition vehicle.

Beremand, D. G.↗
NASA NTRS2014

Concept designs for NASA's Solar Electric Propulsion Technology Demonstration Mission

Multiple Solar Electric Propulsion Technology Demonstration Mission were developed to assess vehicle performance and estimated mission cost. Concepts ranged from a 10,000 kilogram spacecraft capable of delivering 4000 kilogram of payload to one of the Earth Moon Lagrange points in support of future human-crewed outposts to a 180 kilogram spacecraft capable of performing an asteroid rendezvous mission after launched to a geostationary transfer orbit as a secondary payload. Low-cost and maximum Delta-V capability variants of a spacecraft concept based on utilizing a secondary payload adapter as the primary bus structure were developed as were concepts designed to be co-manifested with another spacecraft on a single launch vehicle. Each of the Solar Electric Propulsion Technology Demonstration Mission concepts developed included an estimated spacecraft cost. These data suggest estimated spacecraft costs of $200 million - $300 million if 30 kilowatt-class solar arrays and the corresponding electric propulsion system currently under development are used as the basis for sizing the mission concept regardless of launch vehicle costs. The most affordable mission concept developed based on subscale variants of the advanced solar arrays and electric propulsion technology currently under development by the NASA Space Technology Mission Directorate has an estimated cost of $50M and could provide a Delta-V capability comparable to much larger spacecraft concepts.

solar electric propulsion↗
NASA NTRS2019

Adjoint-Based Design of a Distributed Propulsion Concept with a Power Objective

The adjoint-based design capability in FUN3D is extended to allow efficient gradient-based optimization and design of concepts with highly integrated and distributed aero-propulsive systems. Calculations of propulsive power, along with the derivatives needed to perform adjoint-based design, have been implemented in FUN3D. The design capability is demonstrated by the shape optimization and propulsor sizing of NASA’s PEGASUS aircraft concept. The optimization objective is the minimization of flow power at the aerodynamic interface planes for the wing-mounted propulsors, as well as the tail-cone boundary layer ingestion propulsor, subject to vehicle performance and propulsive constraints.

Ordaz, Irian↗
NASA NTRS2000

Structures Technology for Future Aerospace Systems

An overview of structures technology for future aerospace systems is given. Discussion focuses on developments in component technologies that will improve the vehicle performance, advance the technology exploitation process, and reduce system life-cycle costs. The component technologies described are smart materials and structures, multifunctional materials and structures, affordable composite structures, extreme environment structures, flexible load bearing structures, and computational methods and simulation-based design. The trends in each of the component technologies are discussed and the applicability of these technologies to future aerospace vehicles is described.

Noor, Ahmed K.↗
NASA NTRS1983

Evaluation of g seat augmentation of fixed-base/moving base simulation for transport landings under two visually imposed runway width conditions

Vertical-motion cues supplied by a g-seat to augment platform motion cues in the other five degrees of freedom were evaluated in terms of their effect on objective performance measures obtained during simulated transport landings under visual conditions. In addition to evaluating the effects of the vertical cueing, runway width and magnification effects were investigated. The g-seat was evaluated during fixed base and moving-base operations. Although performance with the g-seat only improved slightly over that with fixed-base operation, combined g-seat platform operation showed no improvement over improvement over platform-only operation. When one runway width at one magnification factor was compared with another width at a different factor, the visual results indicated that the runway width probably had no effect on pilot-vehicle performance. The new performance differences that were detected may be more readily attributed to the extant (existing throughout) increase in vertical velocity induced by the magnification factor used to change the runway width, rather than to the width itself.

Parrish, R. V.↗
NASA NTRS2019

Design Considerations and Development Status for Atmospheric Powered Descent of High-Mass Payloads at Mars

Retropropulsion, initiated at supersonic conditions, is the only presently viable approach to the descent and landing of human-scale payloads on Mars, providing both deceleration and control authority for precision landing. Atmospheric powered descent challenges traditional processes and dependencies in system design. Changes to a design by one discipline have a greater potential to affect overall vehicle performance than in the case of powered descent with no atmosphere. The interdependence is due to the sensitivity of propulsive-aerodynamic interference effects to small changes in configuration or operational environment. The fidelity of Mars exploration architectures has increased in recent years, and the conditions and designs with the lowest aerosciences uncertainties are no longer as flight-relevant. At this time, there is no human-rated vehicle relying on retropropulsion for descent and landing in an atmospheric environment. Continued development of powered descent for atmospheric environments is enabled by progressive balanced investments in ground testing, predictive computational capabilities, and flight testing and demonstration.

EDL↗
NASA NTRS2015

Assimilation of Wind Profiles from Multiple Doppler Radar Wind Profilers for Space Launch Vehicle Applications

Space launch vehicles utilize atmospheric winds in design of the vehicle and during day-of-launch (DOL) operations to assess affects of wind loading on the vehicle and to optimize vehicle performance during ascent. The launch ranges at NASA's Kennedy Space Center co-located with the United States Air Force's (USAF) Eastern Range (ER) at Cape Canaveral Air Force Station and USAF's Western Range (WR) at Vandenberg Air Force Base have extensive networks of in-situ and remote sensing instrumentation to measure atmospheric winds. Each instrument's technique to measure winds has advantages and disadvantages in regards to use for vehicle engineering assessments. Balloons measure wind at all altitudes necessary for vehicle assessments, but two primary disadvantages exist when applying balloon output on DOL. First, balloons need approximately one hour to reach required altitude. For vehicle assessments this occurs at 60 kft (18.3 km). Second, balloons are steered by atmospheric winds down range of the launch site that could significantly differ from those winds along the vehicle ascent trajectory. Figure 1 illustrates the spatial separation of balloon measurements from the surface up to approximately 55 kft (16.8 km) during the Space Shuttle launch on 10 December 2006. The balloon issues are mitigated by use of vertically pointing Doppler Radar Wind Profilers (DRWPs). However, multiple DRWP instruments are required to provide wind data up to 60 kft (18.3 km) for vehicle trajectory assessments. The various DRWP systems have different operating configurations resulting in different temporal and spatial sampling intervals. Therefore, software was developed to combine data from both DRWP-generated profiles into a single profile for use in vehicle trajectory analyses. Details on how data from various wind measurement systems are combined and sample output will be presented in the following sections.

Decker, Ryan K.↗
NASA NTRS2005

Performance Monitoring and Assessment of Neuro-Adaptive Controllers for Aerospace Applications Using a Bayesian Approach

Modem aircraft, UAVs, and robotic spacecraft pose substantial requirements on controllers in the light of ever increasing demands for reusability, affordability, and reliability. The individual systems (which are often nonlinear) must be controlled safely and reliably in environments where it is virtually impossible to analyze-ahead of time- all the important and possible scenarios and environmental factors. For example, system components (e.g., gyros, bearings of reaction wheels, valves) may deteriorate or break during autonomous UAV operation or long-lasting space missions, leading to a sudden, drastic change in vehicle performance. Manual repair or replacement is not an option in such cases. Instead, the system must be able to cope with equipment failure and deterioration. Controllability of the system must be retained as good as possible or re-established as fast as possible with a minimum of deactivation or shutdown of the system being controlled. In such situations the control engineer has to employ adaptive control systems that automatically sense and correct themselves whenever drastic disturbances and/or severe changes in the plant or environment occur.

Gupta, Pramod↗
NASA NTRS2000

GPS Sounding Rocket Development at NASA with Simultaneous Multi-Payload Tracking Application

An inverse differential GPS system has been developed for Sounding Rocket use which includes the flight unit and a ground station capable of extracting GPS data from sounding rocket telemetry, performing a real time differential solution and graphically displaying the rocket's path relative to a predicted trajectory plot. Accuracy has been proven to within less than 10 meters. Postprocessing has increased the precision to within 10 - 20 centimeters. The system has been successfully flown several times and delivered to the Sounding Program Office for routine field use. In addition to providing position, velocity and time GPS data has been used on sounding rockets for vehicle performance analysis, effecting a one hundred fold improvement in data time tagging, and steering an optical tracking device to intercept payloads launched from over the horizon. Precise velocity separation information and timing has been provided to multiple payload systems. Future plans include its use for Range Safety and enabling of interferometric techniques. The technology and software developed also has potential application to small satellite navigation and formation flying.

Bull, Barton↗
NASA NTRS2000

Research Activities within NASA's Morphing Program

In the last decade, smart technologies have become important enabling technologies that cut across traditional boundaries in science and engineering. Here smart is defined as the ability to respond to a stimulus in a predictable and reproducible manner. While multiple successes have been achieved in the laboratory, we have yet to see the general applicability of smart technologies to actual aircraft and spacecraft. The NASA Morphing program is an attempt to couple research across a wide range of disciplines to integrate smart technologies into high payoff applications on aircraft and spacecraft. The program bridges research in several technical disciplines and combines the effort into applications that include active aerodynamic control, active aeroelastic control, and vehicle performance improvement. System studies are used to assess the highest-payoff program objectives, and specific research activities are defined to address the technologies required for development of smart aircraft and spacecraft. This paper will discuss the overall goals of NASA's Morphing program, highlight some of the recent research efforts and discuss the multidisciplinary studies that support that research and some of the challenges associated with bringing the smart technologies to real applications on flight vehicles.

McGowan, Anna-Maria R.↗
NASA NTRS1993

Simulation motion effect on single axis compensatory tracking

An experiment that examined how changes in a motion platform drive filter affect pilot-vehicle performance and opinion was conducted on the NASA Ames Vertical Motion Simulator. Pilots controlled a simplified helicopter model in the vertical or the directional axis and tracked a randomly moving target aircraft in the presence of a random disturbance. With both tasks, variations from full motion to fixed-base conditions were made in the high-pass drive filter gain and natural frequency. The results indicate that vertical motion did not affect the open-loop pilot-vehicle target-tracking crossover frequency, but target-tracking phase margins improved with increased filter gain or decreased natural frequency. Vertical disturbance-rejection crossover frequency increased with decreasing filter natural frequency, while disturbance rejection phase margins improved with increasing filter gain. Vertical tracking errors increased significantly when all vertical motion was removed. No significant differences were measured among the directional configurations, which indicates that pure yaw motion cues may not be as important as previously thought in flight simulation.

Schroeder, Jeffery A.↗
NASA NTRS2019

Hybrid Transportation System Integrated Trajectory Design and Optimization for Mars Landing Site Accessibility

NASA’s Mars Study Capability Team continues the agency’s efforts to study and refine the nation’s plan to field a sustainable human Mars campaign. One of the primary open issues left unanswered during previous analysis cycles was the uncertainty of the impact to the vehicle performance requirement to deliver crew and cargo to the same landing site across multiple mission opportunities. The Mars Study Capability Team has recently developed an integrated trajectory optimization and system closure model to solve the complex interplanetary trajectory optimization using both low-thrust and high-thrust maneuvers. This paper demonstrates the capability of this new integrated trajectory design and optimization method as it applies to the landing site accessibility problem for the Hybrid transportation architecture. The results showed that the current vehicle is capable of reaching up to +20 degree latitude and down to -20 degree latitude in every mission opportunity from 2033 to 2052. However, reaching latitudes beyond +/- 20 required more propellant than the spacecraft is currently designed to carry for many of the mission opportunities.

Chai, Patrick R.↗
NASA NTRS2006

Advanced Ceramics for NASA's Current and Future Needs

Ceramic composites and monolithics are widely recognized by NASA as enabling materials for a variety of aerospace applications. Compared to traditional materials, ceramic materials offer higher specific strength which can enable lighter weight vehicle and engine concepts, increased payloads, and increased operational margins. Additionally, the higher temperature capabilities of these materials allows for increased operating temperatures within the engine and on the vehicle surfaces which can lead to improved engine efficiency and vehicle performance. To meet the requirements of the next generation of both rocket and air-breathing engines, NASA is actively pursuing the development and maturation of a variety of ceramic materials. Anticipated applications for carbide, nitride and oxide-based ceramics will be presented. The current status of these materials and needs for future goals will be outlined. NASA also understands the importance of teaming with other government agencies and industry to optimize these materials and advance them to the level of maturation needed for eventual vehicle and engine demonstrations. A number of successful partnering efforts with NASA and industry will be highlighted.

Jaskowiak, Martha H.↗
NASA NTRS2010

An Up-Close Look at Ares I-X

On October 28, 2009, one day later than the originally planned launch date, the Ares I-X suborbital test flight roared into the Florida sky. Flying its preplanned parabolic arc over the Atlantic, the development test vehicle for the Ares I crew launch vehicle performed as advertised, executing a perfect liftoff, 90-degree roll maneuver, ascent, and separation before its upper and lower stages descended into the ocean 150 miles downrange. This test flight, while carrying no astronauts, marked a major milestone for NASA, which had not flown a test launch of a human-rated rocket since the first flight of the Space Shuttle in 1981. During the flight, over 700 sensors collected over 900 measurements, which NASA will apply to validating the engineering models they used to design the vehicle in the first place. That data, telemetered to the ground and stored in a flight recorder onboard, was the primary "payload" of the mission.

Leahy, Bart↗