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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 217 records · Page 12
All sourcesNASA NTRSDOE OSTI
NASA NTRS2010

Evaluation of Dual-Launch Lunar Architectures Using the Mission Assessment Post Processor

The National Aeronautics and Space Administrations (NASA) Constellation Program is currently designing a new transportation system to replace the Space Shuttle, support human missions to both the International Space Station (ISS) and the Moon, and enable the eventual establishment of an outpost on the lunar surface. The present Constellation architecture is designed to meet nominal capability requirements and provide flexibility sufficient for handling a host of contingency scenarios including (but not limited to) launch delays at the Earth. This report summarizes a body of work performed in support of the Review of U.S. Human Space Flight Committee. It analyzes three lunar orbit rendezvous dual-launch architecture options which incorporate differing methodologies for mitigating the effects of launch delays at the Earth. NASA employed the recently-developed Mission Assessment Post Processor (MAPP) tool to quickly evaluate vehicle performance requirements for several candidate approaches for conducting human missions to the Moon. The MAPP tool enabled analysis of Earth perturbation effects and Earth-Moon geometry effects on the integrated vehicle performance as it varies over the 18.6-year lunar nodal cycle. Results are provided summarizing best-case and worst-case vehicle propellant requirements for each architecture option. Additionally, the associated vehicle payload mass requirements at launch are compared between each architecture and against those of the Constellation Program. The current Constellation Program architecture assumes that the Altair lunar lander and Earth Departure Stage (EDS) vehicles are launched on a heavy lift launch vehicle. The Orion Crew Exploration Vehicle (CEV) is separately launched on a smaller man-rated vehicle. This strategy relaxes man-rating requirements for the heavy lift launch vehicle and has the potential to significantly reduce the cost of the overall architecture over the operational lifetime of the program. The crew launch occurs first, four days prior to the optimal trans-lunar injection (TLI) departure window. This is done to allow for launch delays in the Altair/EDS launch. During this time, the Orion vehicle is required to conduct orbit maintenance while loitering in low Earth orbit (LEO). The alternative architectures presented aim to eliminate the need for costly orbit maintenance maneuvers while loitering in LEO. In all of the alternative architectures considered, it is assumed that the Altair and Orion vehicles are nominally launched 90 minutes apart, depart the Earth separately, and complete the rendezvous and docking sequence at the Moon. In this lunar orbit rendezvous (LOR) strategy, both the Altair and Orion vehicles will require separate EDS stages, and each will be required to perform lunar orbit insertion (LOI). This has the effect of balancing payload requirements between the two launch vehicles at the Earth. In this case, the overall payload mass is increased slightly, but the increased mission costs could potentially be offset by requiring the construction of two rockets similar in size and nature, unlike the current Constellation architecture. Three dual-launch architecture options with LOR were evaluated, which incorporate differing methodologies for mitigating the effects of launch delays at the Earth. Benefits and drawbacks of each of the dual-launch architecture options with LOR are discussed and the overall mission performance is compared with that of the existing Constellation Program lunar architecture.

Stewart, Shaun M.↗
NASA NTRS2014

Space Shuttle Launch Probability Analysis: Understanding History so We Can Predict the Future

The Space Shuttle was launched 135 times and nearly half of those launches required 2 or more launch attempts. The Space Shuttle launch countdown historical data of 250 launch attempts provides a wealth of data that is important to analyze for strictly historical purposes as well as for use in predicting future launch vehicle launch countdown performance. This paper provides a statistical analysis of all Space Shuttle launch attempts including the empirical probability of launch on any given attempt and the cumulative probability of launch relative to the planned launch date at the start of the initial launch countdown. This information can be used to facilitate launch probability predictions of future launch vehicles such as NASA's Space Shuttle derived SLS. Understanding the cumulative probability of launch is particularly important for missions to Mars since the launch opportunities are relatively short in duration and one must wait for 2 years before a subsequent attempt can begin.

Cates, Grant R.↗
NASA NTRS2006

The Role of Flight Experiments in the Development of Cryogenic Fluid Management Technologies

This paper reviews the history of cryogenic fluid management technology development and infusion into both the Saturn and Centaur vehicles. Ground testing and analysis proved inadequate to demonstrate full scale performance. As a consequence flight demonstration with a full scale vehicle was required by both the Saturn and Centaur programs to build confidence that problems were addressed. However; the flight vehicles were highly limited on flight instrumentation and the flight demonstration locked-in the design without challenging the function of design elements. Projects reviewed include: the Aerobee Sounding Rocket Cryogenic Fluid Management (CFM) tests which served as a valuable stepping stone to flight demonstration and built confidence in the ability to handle hydrogen in low gravity; the Saturn IVB Fluid Management Qualification flight test; the Atlas Centaur demonstration flights to develop two burn capability; and finally the Titan Centaur two post mission flight tests.

Chato, David J.↗
NASA NTRSDate not supplied

NASA Reference Motor Designs for Electric Vertical Takeoff and Landing Vehicles

Electric and hybrid electric vertical takeoff and landing vehicles require high performance and high reliability electric motor drivetrains. Failure analysis of NASA’s Revolutionary Vertical Lift Technologies’ reference vehicles pointed to current electric motor drivetrain reliability being below what is needed to meet the expected stringent reliability requirements for Urban Air Mobility vehicles. In this paper, design studies are carried out for UAM vehicle electric motors to produce reference designs. The primary intent of these reference motor designs is to provide guidance for UAM motor reliability model development and technology advancement. They additionally provide high fidelity motor sizing information for vehicle designers and references for different technologies or motor topologies to be traded against.

Electric Motor Urban Air Mobility↗
NASA NTRSDate not supplied

NASA Reference Motor Designs for Electric Vertical Takeoff and Landing Vehicles

Electric and hybrid electric vertical takeoff and landing vehicles require high performance and high reliability electric motor drivetrains. Failure analysis of NASA’s Revolutionary Vertical Lift Technologies’ reference vehicles pointed to current electric motor drivetrain reliability being below what is needed to meet the expected stringent reliability requirements for Urban Air Mobility vehicles [1]. In this paper, design studies are carried out for UAM vehicle electric motors to produce reference designs. The primary intent of these reference motor designs is to provide guidance for UAM motor reliability model development and technology advancement. They additionally provide high fidelity motor sizing information for vehicle designers and references for different technologies or motor topologies to be traded against.

Thomas Tallerico↗
NASA NTRS1990

Active control of aerothermoelastic effects for a conceptual hypersonic aircraft

Procedures for and results of aeroservothermoelastic studies are described. The objectives of these studies were to develop the necessary procedures for performing an aeroelastic analysis of an aerodynamically heated vehicle and to analyze a configuration in the classical cold state and in a hot state. Major tasks include the development of the structural and aerodynamic models, open loop analyses, design of active control laws for improving dynamic responses and analyses of the closed loop vehicles. The analyses performed focused on flutter speed calculations, short period eigenvalue trends and statistical analyses of the vehicle response to controls and turbulence. Improving the ride quality of the vehicle and raising the flutter boundary of the aerodynamically-heated vehicle up to that of the cold vehicle were the objectives of the control law design investigations.

Heeg, Jennifer↗
NASA NTRS1990

Active control of aerothermoelastic effects for a conceptual hypersonic aircraft

This paper describes the procedures for an results of aeroservothermoelastic studies. The objectives of these studies were to develop the necessary procedures for performing an aeroelastic analysis of an aerodynamically heated vehicle and to analyze a configuration in the classical 'cold' state and in a 'hot' state. Major tasks include the development of the structural and aerodynamic models, open loop analyses, design of active control laws for improving dynamic responses and analyses of the closed loop vehicles. The analyses performed focused on flutter speed calculations, short period eigenvalue trends and statistical analyses of the vehicle response to controls and turbulence. Improving the ride quality of the vehicle and raising the flutter boundary of the aerodynamically-heated vehicle up to that of the cold vehicle were the objectives of the control law design investigations.

Heeg, Jennifer↗
NASA NTRS2020

Pterodactyl: The Development and Performance of Guidance Algorithms for a Mechanically Deployed Entry Vehicle

Pterodactyl is a NASA Space Technology Mission Directorate (STMD) project focused on developing a design capability for optimal, scalable, Guidance and Control (G&C) solutions that enable precision targeting for Deployable Entry Vehicles (DEVs). This feasibility study is unique in that it focuses on the rapid integration of targeting performance analysis with structural & packaging analysis, which is especially challenging for new vehicle and mission designs. This paper will detail the guidance development and trajectory design process for a lunar return mission, selected to stress the vehicle designs and encourage future scalability. For the five G&C configurations considered, the Fully Numerical Predictor-Corrector Entry Guidance (FNPEG) was selected for configurations requiring bank angle guidance and FNPEG with Uncoupled Range Control (URC) was developed for configurations requiring angle of attack and sideslip angle guidance. Successful G&C configurations are defined as those that can deliver payloads to the intended descent and landing initiation point, while abiding by trajectory constraints for nominal and dispersed trajectories.

Johnson, Breanna J.↗
NASA NTRS1986

Thrust vectoring for single-stage-to-orbit, horizontal takeoff, horizontal landing, space vehicles

The preliminary design of a horizontal takeoff, horizontal landing, single-stage-to-orbit, rocket-powered space vehicle was performed. The purpose was to examine technology requirements for future small payload launch vehicles. The distinguishing aspect of the design was the utilization of thrust vectoring to provide half of the lift at takeoff. The inclusion of a canard was necessary to provide additional lift at takeoff and to balance the moments produced with thrust vectoring. A weights estimation, an aerodynamic assessment, a trajectory analysis, and a gear weight analysis were performed. The takeoff weight of the resulting vehicle was approximately 1.26 million pounds, based on advanced technology structures and subsystems. The vehicle was designed to deliver a 5000-pound payload to a polar orbit.

Cunningham, M. J.↗
NASA NTRS2013

Launch Vehicle Abort Analysis for Failures Leading to Loss of Control

Launch vehicle ascent is a time of high risk for an onboard crew. There is a large fraction of possible failures for which time is of the essence and a successful abort is possible if the detection and action happens quickly enough. This paper focuses on abort determination based on data already available from the Guidance, Navigation, and Control system. This work is the result of failure analysis efforts performed during the Ares I launch vehicle development program. The two primary areas of focus are the derivation of abort triggers to ensure that abort occurs as quickly as possible when needed, but that false aborts are avoided, and evaluation of success in aborting off the failing launch vehicle.

Hanson, John M.↗
NASA NTRS1975

Investigation of lightweight designs and materials for LO2 and LH2 propellant tanks for space vehicles, phase 1

Design, analysis, and fabrication studies were performed on nonintegral (suspended) tanks using a representative space tug design. The LH2 and LO2 tank concept selection was developed. Tank geometries and support relationships were investigated using tug design propellant inertias and ullage pressures, then compared based on total tug systems effects. The tank combinations which resulted in the maximum payload were selected. Tests were conducted on samples of membrane material which was processed in a manner simulating production tank fabrication operations to determine fabrication effects on the fracture toughness of the tank material. Fracture mechanics analyses were also performed to establish a preliminary set of allowables for initial defects.

Source record↗
NASA NTRS1986

Evaluation of a Stirling engine heater bypass with the NASA Lewis nodal-analysis performance code

In support of the U.S. Department of Energy's Stirling Engine Highway Vehicle Systems program, the NASA Lewis Research Center investigated whether bypassing the P-40 Stirling engine heater during regenerative cooling would improve engine performance. The Lewis nodal-analysis Stirling engine computer simulation was used for this investigation. Results for the heater-bypass concept showed no significant improvement in the indicated thermal efficiency for the P-40 Stirling engine operating at full-power and part-power conditions. Optimizing the heater tube length produced a small increase in the indicated thermal efficiency with the heater-bypass concept.

Sullivan, T. J.↗
NASA NTRS1991

Evaluation of on-board hydrogen storage methods f or high-speed aircraft

Hydrogen is the fuel of choice for hypersonic vehicles. Its main disadvantage is its low liquid and solid density. This increases the vehicle volume and hence the drag losses during atmospheric flight. In addition, the dry mass of the vehicle is larger due to larger vehicle structure and fuel tankage. Therefore it is very desirable to find a fuel system with smaller fuel storage requirements without deteriorating the vehicle performance substantially. To evaluate various candidate fuel systems, they were first screened thermodynamically with respect to their energy content and cooling capacities. To evaluate the vehicle performance with different fuel systems, a simple computer model is developed to compute the vehicle parameters such as the vehicle volume, dry mass, effective specific impulse, and payload capacity. The results indicate that if the payload capacity (or the gross lift-off mass) is the most important criterion, only slush hydrogen and liquid hydrogen - liquid methane gel shows better performance than the liquid hydrogen vehicle. If all the advantages of a smaller vehicle are considered and a more accurate mass analysis can be performed, other systems using endothermic fuels such as cyclohexane, and some boranes may prove to be worthy of further consideration.

Akyurtlu, Ates↗
NASA NTRS1994

A parametric sensitivity study for single-stage-to-orbit hypersonic vehicles using trajectory optimization

The class of hypersonic vehicle configurations with single stage-to-orbit (SSTO) capability reflect highly integrated airframe and propulsion systems. These designs are also known to exhibit a large degree of interaction between the airframe and engine dynamics. Consequently, even simplified hypersonic models are characterized by tightly coupled nonlinear equations of motion. In addition, hypersonic SSTO vehicles present a major system design challenge; the vehicle's overall mission performance is a function of its subsystem efficiencies including structural, aerodynamic, propulsive, and operational. Further, all subsystem efficiencies are interrelated, hence, independent optimization of the subsystems is not likely to lead to an optimum design. Thus, it is desired to know the effect of various subsystem efficiencies on overall mission performance. For the purposes of this analysis, mission performance will be measured in terms of the payload weight inserted into orbit. In this report, a trajectory optimization problem is formulated for a generic hypersonic lifting body for a specified orbit-injection mission. A solution method is outlined, and results are detailed for the generic vehicle, referred to as the baseline model. After evaluating the performance of the baseline model, a sensitivity study is presented to determine the effect of various subsystem efficiencies on mission performance. This consists of performing a parametric analysis of the basic design parameters, generating a matrix of configurations, and determining the mission performance of each configuration. Also, the performance loss due to constraining the total head load experienced by the vehicle is evaluated. The key results from this analysis include the formulation of the sizing problem for this vehicle class using trajectory optimization, characteristics of the optimal trajectories, and the subsystem design sensitivities.

Lovell, T. Alan↗
NASA NTRS1977

Updated dispersion analysis for the first Orbital Flight Test (OFT-1) mission

A dispersion analysis considering 3-sigma uncertainties (or perturbations) in platform, vehicle, and environmental parameters was performed for the first orbital flight test (OFT-1) mission. The dispersion analysis is based on the nominal trajectory for the OFT-1 reference flight profile and was performed to determine state vector and performance dispersions (or variations) which result from the indicated 3 sigma uncertainties. The dispersions are determined at major mission events and fixed times from liftoff (time slices). Principal error contributors to the covariance matrix are listed. The dispersion data indicates that the largest position error occurs in the down range component. At main engine cutoff and circularization, the vehicle performance uncertainties are the major contributors to down range error.

Snow, L. S.↗
NASA NTRS1993

Aerothermoelastic analysis of a NASP demonstrator model

The proposed National AeroSpace Plane (NASP) is designed to travel at speeds up to Mach 25. Because aerodynamic heating during high-speed flight through the atmosphere could destiffen a structure, significant couplings between the elastic and rigid body modes could result in lower flutter speeds and more pronounced aeroelastic response characteristics. These speeds will also generate thermal loads on the structure. The purpose of this research is develop methodologies applicable to the NASP and to apply them to a representative model to determine its aerothermoelastic characteristics when subjected to these thermal loads. This paper describes an aerothermoelastic analysis of the generic hypersonic vehicle configuration. The steps involved in this analysis were: (1) generating vehicle surface temperatures at the appropriate flight conditions; (2) applying these temperatures to the vehicle's structure to predict changes in the stiffness resulting from material property degradation; (3) predicting the vibration characteristics of the heated structure at the various temperature conditions; (4) performing aerodynamic analyses; and (5) conducting flutter analysis of the heated vehicle. Results of these analyses and conclusions representative of a NASP vehicle are provided in this paper.

Heeg, Jennifer↗