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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 325 records · Page 18
All sourcesNASA NTRSDOE OSTI
NASA NTRS1979

Flight design system level C requirements. Solid rocket booster and external tank impact prediction processors

The prediction of the SRB and ET impact areas requires six separate processors. The SRB impact prediction processor computes the impact areas and related trajectory data for each SRB element. Output from this processor is stored on a secure file accessible by the SRB impact plot processor which generates the required plots. Similarly the ET RTLS impact prediction processor and the ET RTLS impact plot processor generates the ET impact footprints for return-to-launch-site (RTLS) profiles. The ET nominal/AOA/ATO impact prediction processor and the ET nominal/AOA/ATO impact plot processor generate the ET impact footprints for non-RTLS profiles. The SRB and ET impact processors compute the size and shape of the impact footprints by tabular lookup in a stored footprint dispersion data base. The location of each footprint is determined by simulating a reference trajectory and computing the reference impact point location. To insure consistency among all flight design system (FDS) users, much input required by these processors will be obtained from the FDS master data base.

Seale, R. H.↗
NASA NTRS1994

Design of a Flush Airdata System (FADS) for the Hypersonic Air Launched Option (HALO) Vehicle

This paper presents a design study for a pressure based Flush airdata system (FADS) on the Hypersonic Air Launched Option (HALO) Vehicle. The analysis will demonstrate the feasibility of using a pressure based airdata system for the HALO and provide measurement uncertainty estimates along a candidate trajectory. The HALO is a conceived as a man-rated vehicle to be air launched from an SR-71 platform and is proposed as a testbed for an airbreathing hydrogen scramjet. A feasibility study has been performed and indicates that the proposed trajectory is possible with minimal modifications to the existing SR71 vehicle. The mission consists of launching the HALO off the top of an SR-71 at Mach 3 and 80,000 ft. A rocket motor is then used to accelerate the vehicle to the test condition. After the scramjet test is completed the vehicle will glide to a lakebed runway landing. This option provides reusability of the vehicle and scramjet engine. The HALO design will also allow for various scramjet engine and flowpath designs to be flight tested. For the HALO flights, measurements of freestream airdata are considered to be a mission critical to perform gain scheduling and trajectory optimization. One approach taken to obtaining airdata involves measurement of certain parameters such as external atmospheric winds, temperature, etc to estimate the airdata quantities. This study takes an alternate approach. Here the feasibility of obtaining airdata using a pressure-based flush airdata system (FADS) methods is assessed. The analysis, although it is performed using the HALO configuration and trajectory, is generally applicable to other hypersonic vehicles. The method to be presented offers the distinct advantage of inferring total pressure, Mach number, and flow incidence angles, without stagnating the freestream flow. This approach allows for airdata measurements to be made using blunt surfaces and significantly diminishes the heating load at the sensor. In the FADS concept a matrix of flush ports is placed in the vicinity of the aircraft nose, and the airdata are inferred indirectly from the measured pressures.

Whitmore, Stephen A.↗
NASA NTRS1984

The plasma wave environment of an auroral arc - Electrostatic ion cyclotron waves in the diffuse aurora

Electric field plasma wave observations were made with a sounding rocket payload in and near a quiet auroral arc. This payload was launched on March 9, 1978 from Poker Flat, Alaska. The payload trajectory was close to the magnetic meridian and passed over a 40 kR auroral arc. The present investigation is concerned with ac electric field observations in the ELF and lower VLF covering a frequency range from 2.5 Hz to 8 kHz. Attention is given to aspects of instrumentation, the general situation, a data analysis, and the obtained results. Waves were observed at low altitude in a region of downward parallel current and diffuse aurora. These waves had properties consistent with those expected for hydrogen and oxygen electrostatic ion cyclotron (EIC) waves.

Bering, E. A.↗
NASA NTRS1986

A guidance-motivated sensitivity analysis of an aero-assisted boost vehicle

A simple model of an aero-assisted booster is used to examine the contributions of propulsion system type, aerodynamic lift and flight trajectory to the efficiency with which payloads can be placed into low earth orbit. The higher propulsive efficiency of ramjet and scramjet propulsion has the potential of increasing the payload mass ratio significantly. The contributions of turbojet propulsion and aerodynamic lift are less significant. The additional complexity involved in using aerodynamic propulsion and lift requires dealing with a more comprehensive set of design variables than for rocket boosters. The approach taken is to derive a set of sensitivity functions which relate booster performance to the design variables. The problems of optimum mixing of aerodynamic lift with thrust and determining the optimal boost trajectory is treated. The potential payload capacity of a horizontal take-off air-breathing boost vehicle is examined. The optimization problem which considers propulsive efficiency, aerodynamic configuration, and control and guidance issues is discussed.

Taylor, L. W.↗
NASA NTRS2003

Launch Condition Deviations of Reusable Launch Vehicle Simulations in Exo-Atmospheric Zoom Climbs

The Defense Advanced Research Projects Agency has proposed a two-stage system to deliver a small payload to orbit. The proposal calls for an airplane to perform an exo-atmospheric zoom climb maneuver, from which a second-stage rocket is launched carrying the payload into orbit. The NASA Dryden Flight Research Center has conducted an in-house generic simulation study to determine how accurately a human-piloted airplane can deliver a second-stage rocket to a desired exo-atmospheric launch condition. A high-performance, fighter-type, fixed-base, real-time, pilot-in-the-loop airplane simulation has been modified to perform exo-atmospheric zoom climb maneuvers. Four research pilots tracked a reference trajectory in the presence of winds, initial offsets, and degraded engine thrust to a second-stage launch condition. These launch conditions have been compared to the reference launch condition to characterize the expected deviation. At each launch condition, a speed change was applied to the second-stage rocket to insert the payload onto a transfer orbit to the desired operational orbit. The most sensitive of the test cases was the degraded thrust case, yielding second-stage launch energies that were too low to achieve the radius of the desired operational orbit. The handling qualities of the airplane, as a first-stage vehicle, have also been investigated.

Urschel, Peter H.↗
NASA NTRS1976

An experimental determination in Calspan Ludwieg tube of the base environment of the integrated space shuttle vehicle at simulated Mach 4.5 flight conditions (test IH5 of model 19-OTS)

A model test program was conducted to determine heat transfer and pressure distributions in the base region of the space shuttle vehicle during simulated launch trajectory conditions of Mach 4.5 and pressure altitudes between 90,000 and 210,000 feet. Model configurations with and without the solid propellant booster rockets were examined to duplicate pre- and post-staging vehicle geometries. Using short duration flow techniques, a tube wind tunnel provided supersonic flow over the model. Simultaneously, combustion generated exhaust products reproduced the gasdynamic and thermochemical structure of the main vehicle engine plumes. Heat transfer and pressure measurements were made at numerous locations on the base surfaces of the 19-OTS space shuttle model with high response instrumentation. In addition, measurements of base recovery temperature were made indirectly by using dual fine wire and resistance thermometers and by extrapolating heat transfer measurements.

Drzewiecki, R. F.↗
NASA NTRS2011

COMPASS Final Report: Radioisotope Electric Propulsion (REP) Centaur Orbiter New Frontiers Mission

Radioisotope Electric Propulsion (REP) has been shown in past studies to enable missions to outer planetary bodies including the orbiting of Centaur asteroids. Key to the feasibility for REP missions are long life, low power electric propulsion (EP) devices, low mass Radioisotope Power System (RPS) and light spacecraft (S/C) components. In order to determine the key parameters for EP devices to perform these REP missions a design study was completed to design an REP S/C to orbit a Centaur in a New Frontiers (NF) cost cap. The design shows that an orbiter using several long lived (approx.200 kg xenon (Xe) throughput), low power (approx.700 W) Hall thrusters teamed with six (150 W each) Advanced Stirling Radioisotope Generators (ASRG) can deliver 60 kg of science instruments to a Centaur in 10 yr within the NF cost cap. Optimal specific impulses (Isp) for the Hall thrusters were found to be around 2000 s with thruster efficiencies over 40 percent. Not only can the REP S/C enable orbiting a Centaur (when compared to an all chemical mission only capable of flybys) but the additional power from the REP system can be used to enhance science and simplify communications. The mission design detailed in this report is a Radioisotope Power System (RPS) powered EP science orbiter to the Centaur Thereus with arrival 10 yr after launch, ending in a 1 yr science mapping mission. Along the trajectory, approximately 1.5 yr into the mission, the REP S/C does a flyby of the Trojan asteroid Tlepolemus. The total (Delta)V of the trajectory is 8.9 km/s. The REP S/C is delivered to orbit on an Atlas 551 class launch vehicle with a Star 48 B solid rocket stage

Oleson, Steven R.↗
NASA NTRSDate not supplied

ACS3: Flight Dynamics

The NASA's Advanced Composite Solar Sail System (ACS3) mission consist of a spacecraft that will deploy an 80 m2 solar sail in Low Earth Orbit (LEO). The main objective of the mission is to demonstrate that the solar wind can impulse the spacecraft to change the semimajor-axis and obtain a different orbit altitude. The sail will be composed of a combination of composite materials with distinct properties, and it will be deployed with lightweight booms from a 12U CubeSat bus, developed by Nanoavionics. The spacecraft will be launched aboard an Electron launch vehicle from Rocket LAB Launch Complex in New Zealand. This paper covers the orbital mechanics and navigation developments to support the mission, from the solar sail trajectory model to the actual flight dynamics system to provide the orbit determination analysis prior to flight. First, we introduce a description of our high-fidelity propagation that accounts for the solar radiation pressure to produce predictive ephemeris of the solar sail performance with several spacecraft attitude modes. As part of our results, we present plots of the expected altitudes achieved by the spacecraft once the solar sail is deployed under various assumptions of the solar weather. In addition to that, we present a full description of our orbit determination process which relies in GPS state vectors to accurately estimate the position and velocity uncertainties at a frequent cadence during the mission. The outcome of this process will be critical to achieving the objective of determining effective altitude change produced by the solar sail.

Solar sail↗
NASA NTRS1984

Ion resonances and ELF wave production by an electron beam injected into the ionosphere - ECHO 6

Two effects observed with electron antennas ejected from a sounding rocket launched into the ionosphere in March 1983 carrying electron beam guns are discussed. The sensor packages were ejected and travelled parallel to the vehicle trajectory. Electric potentials were measured between the single probes and a plasma diagnostic package while the gun injected electrons into the ionosphere in perpendicular and parallel 1 kHz directions. Signal pulses over the dc-1250 kHz range were detected. A kHz gun frequency caused a signal that decreased by two orders of magnitude between 45-90 m from the beam field line. However, the signal was detectable at 1 mV/m at 120 m, supporting earlier data that indicated that pulsed electron beams can cause ELF waves in space. Beam injection parallel to the magnetic field produced an 840 Hz resonance that could be quenched by activation of a transverse beam.

Winckler, J. R.↗
NASA NTRS1986

STS performance enhancements for a heads up ascent design

A method of increasing the STS (space transportation system) lift capability by flying the SSV (Space Shuttle vehicle) in a heads up attitude during the first stage atmospheric flight as opposed to the current heads down attitude is presented. The paper reveals that the STS lift capability and SRM (solid rocket motor) burnrate definition are sensitive to the magnitude and direction of normal aerodynamic force acting on the SSV during atmospheric ascent. The STS operational trajectory design angle of attack will permit increases in SRM burnrate definition and STS lift capability for the proposed heads up ascent when compared to the current heads down ascent. For a heads up ascent the SRM burnrate can be increased 7 percent with a corresponding 4000 lb increase in STS lift capability. A comparison of the heads up and heads down ascent reveals similar performance trades for changes in trajectory shaping parameters and SRM and SSME (Space Shuttle main engine) performance characteristics.

Millard, S. W.↗
NASA NTRS1993

Prospects for utilization of air liquefaction and enrichment system (ALES) propulsion in fully reusable launch vehicles

A concept is shown for a fully reusable, earth to orbit launch vehicle with horizontal takeoff and landing, employing an air-turborocket for low speed and a rocket for high speed acceleration, both using LH2 fuel. The turborocket employs a modified liquid air cycle to supply the oxidizer. The rocket uses 90 percent pure LOX that is collected from the atmosphere, separated, and stored during operation of the turborocket from about Mach 2 to Mach 5 or 6. The takeoff weight and the thrust required at takeoff are markedly reduced by collecting the rocket oxidizer in-flight. The paper shows an approach and the corresponding technology needs for using ALES propulsion in a SSTO vehicle. Reducing the trajectory altitude at the end of collection reduces the wing area and increases payload. The use of state-of-the-art materials, such as graphite polyimide, is critical to meet the structure weight objective for SSTO. Configurations that utilize 'waverider' aerodynamics show great promise to reduce the vehicle weight.

Bond, W. H.↗
NASA NTRS2011

Ares I-X Test Flight Reference Trajectory Development

Ares I-X was the first test flight of NASA's Constellation Program's Ares I crew launch vehicle. Ares I is a two stage to orbit launch vehicle that provides crew access to low Earth orbit for NASA's future manned exploration missions. The Ares I first stage consists of a Shuttle solid rocket motor (SRM) modified to include an additional propellant segment and a liquid propellant upper stage with an Apollo J2X engine modified to increase its thrust capability. The modified propulsion systems were not available for the first test flight, thus the test had to be conducted with an existing Shuttle 4 segment reusable solid rocket motor (RSRM) and an inert Upper Stage. The test flight's primary objective was to demonstrate controllability of an Ares I vehicle during first stage boost and the ability to perform a successful separation. In order to demonstrate controllability, the Ares I-X ascent control algorithms had to maintain stable flight throughout a flight environment equivalent to Ares I. The goal of the test flight reference trajectory development was to design a boost trajectory using the existing RSRM that results in a flight environment equivalent to Ares I. A trajectory similarity metric was defined as the integrated difference between the Ares I and Ares I-X Mach versus dynamic pressure relationships. Optimization analyses were performed that minimized the metric by adjusting the inert upper stage weight and the ascent steering profile. The sensitivity of the optimal upper stage weight and steering profile to launch month was also investigated. A response surface approach was used to verify the optimization results. The analyses successfully defined monthly ascent trajectories that matched the Ares I reference trajectory dynamic pressure versus Mach number relationship to within 10% through Mach 3.5. The upper stage weight required to achieve the match was found to be feasible and varied less than 5% throughout the year. The paper will discuss the flight test requirements, provide Ares I-X vehicle background, discuss the optimization analyses used to meet the requirements, present analysis results, and compare the reference trajectory to the reconstructed flight trajectory.

Starr, Brett R.↗
NASA NTRS1992

A two-phase restricted equilibrium model for combustion of metalized solid propellants

An Eulerian-Lagrangian two-phase approach was adopted to model the multi-phase reacting internal flow in a solid rocket with a metalized propellant. An Eulerian description was used to analyze the motion of the continuous phase which includes the gas as well as the small (micron-sized) particulates, while a Lagrangian description is used for the analysis of the discrete phase which consists of the larger particulates in the motor chamber. The particulates consist of Al and Al2O3 such that the particulate composition is 100 percent Al at injection from the propellant surface with Al2O3 fraction increasing due to combustion along the particle trajectory. An empirical model is used to compute the combustion rate for agglomerates while the continuous phase chemistry is treated using chemical equilibrium. The computer code was used to simulate the reacting flow in a solid rocket motor with an AP/HTPB/Al propellant. The computed results show the existence of an extended combustion zone in the chamber rather than a thin reaction region. The presence of the extended combustion zone results in the chamber flow field and chemical being far from isothermal (as would be predicted by a surface combustion assumption). The temperature in the chamber increases from about 2600 K at the propellant surface to about 3350 K in the core. Similarly the chemical composition and the density of the propellant gas also show spatially non-uniform distribution in the chamber. The analysis developed under the present effort provides a more sophisticated tool for solid rocket internal flow predictions than is presently available, and can be useful in studying apparent anomalies and improving the simple correlations currently in use. The code can be used in the analysis of combustion efficiency, thermal load in the internal insulation, plume radiation, etc.

Sabnis, J. S.↗
NASA NTRS2010

Contingency Operations of Americas Next Moon Rocket, Ares V

America has begun the development of a new space vehicle system which will enable humans to return to the moon and reach even farther destinations. The system is called Constellation: it has 2 earth-launch vehicles, Ares I and Ares V; a crew module, Orion; and a lander, Altair with descent and ascent stages. Ares V will launch an Earth Departure Stage (EDS) and Altair into low earth orbit. Ares I will launch the Orion crew module into low earth orbit where it will rendezvous and dock with the Altair and EDS "stack". After rendezvous, the stack will contain four complete rocket systems, each capable of independent operations. Of course this multiplicity of vehicles provides a multiplicity of opportunities for off-nominal behavior and multiple mitigation options for each. Contingency operations are complicated by the issues of crew safety and the possibility of debris from the very large components impacting the ground. This paper examines contingency operations of the EDS in low earth orbit, during the boost to translunar orbit, and after the translunar boost. Contingency operations under these conditions have not been a consideration since the Apollo era and analysis of the possible contingencies and mitigations will take some time to evolve. Since the vehicle has not been designed, much less built, it is not possible to evaluate contingencies from a root-cause basis or from a probability basis; rather they are discussed at an effects level (such as the reaction control system is consuming propellant at a high rate). Mitigations for the contingencies are based on the severity of the off-nominal condition, the time of occurrence, recovery options, options for alternate missions, crew safety, evaluation of the condition (forensics) and future prevention. Some proposed mitigations reflect innovation in thinking and make use of the multiplicity of on-orbit resources including the crew; example: Orion could do a "fly around" to allow the crew to determine the condition and cause of a partially separated payload shroud. Other mitigations are really alternate missions; example, an engine out on during ascent resulted in insufficient propellant for the lunar mission, but the on-orbit vehicle stack is otherwise perfect and can pursue an alternate mission, such as a high ballistic trajectory to test the high-speed atmospheric reentry of Orion. Evaluation and presentation of contingency operations at this early stage of the development of the Ares V rocket will improve the design of the vehicle and lay the groundwork for the exhaustive contingency planning which must be done after the vehicle is built as preparations for operations.

Jaap, John↗
NASA NTRS2017

Development of Models for Disk-Gap-Band Parachutes Deployed Supersonically in the Wake of a Slender Body

The Advanced Supersonic Parachute Inflation Research and Experiments (ASPIRE) project will investigate the supersonic deployment, inflation, and aerodynamics of Disk- Gap-Band (DGB) parachutes in the wake of a slender body. The parachutes will be full-scale versions of the DGBs used by the Mars Science Laboratory in 2012 and planned for NASA’s Mars 2020 project and will be delivered to targeted deployment conditions representative of flight at Mars by sounding rockets launched out of NASA’s Wallops Flight Facility. The parachutes will be tested in the wake of a slender payload whose diameter is approximately a sixth that of entry capsules used for Mars missions. Models of the deployment, inflation, and aerodynamic performance of the parachutes are necessary to design key aspects of the experiment, including: determining the expected loads and applicable margins on the parachute and payload; guiding sensor selection and placement; evaluating the vehicle trajectory for targeting, range safety, and recovery operations. In addition, knowledge of the differences in the behavior of the parachutes in the wake of slender and blunt bodies is required in order to interpret the results of the sounding rocket experiment and determine how they relate to expected performance behind blunt bodies at Mars. However, modeling the performance of a supersonic DGB in the wake of a slender body is challenging due to the scarcity of historical test data and modeling precedents. This paper describes the models of the aerodynamic performance of DGBs in supersonic slender-body wakes being developed for the ASPIRE sounding rocket test campaign. Development of these models is based on the four available flight tests of DGBs deployed in supersonic slender-body wakes as well as on data from past flight and wind-tunnel experiments of DGBs deployed in the wake of blunt bodies, on the reconstructed at-Mars DGB performance during past missions, and on computational fluid dynamics simulations. Simulations of the wakes of blunt and slender bodies in supersonic flow have been conducted in order to investigate the differences in the flowfields encountered by parachutes deployed in both wake types. The simulations have allowed the project to investigate the fundamental differences between the sounding rocket tests and the flight of a DGB during a Mars mission and to assess the limitations of the sounding rocket test architecture for evaluating flight performance at Mars.

Clark, Ian G.↗
NASA NTRS1976

Aerodynamic results of wind tunnel tests on a 0.010-scale model (32-QTS) space shuttle integrated vehicle in the AEDC VKF-40-inch supersonic wind tunnel (IA61)

Plotted and tabulated aerodynamic coefficient data from a wind tunnel test of the integrated space shuttle vehicle are presented. The primary test objective was to determine proximity force and moment data for the orbiter/external tank and solid rocket booster (SRB) with and without separation rockets firing for both single and dual booster runs. Data were obtained at three points (t = 0, 1.25, and 2.0 seconds) on the nominal SRB separation trajectory.

Daileda, J. J.↗
NASA NTRS2010

ARES I-X: The First Test Flight of a New Era

On October 28th, 2009, the National Aeronautics and Space Administration (NASA) launched the Ares I-X Flight Test Vehicle (FTV) from pad 39B, providing the first set of flight test data for NASA's Ares I vehicle design team. This test was critical in providing insight into areas were significant design challenges existed. This paper discusses the objectives of the mission and how they were satisfied. It discusses the overall results of the flight test and look at the data retrieved from the flight. Ares I-X was highly instrumented with over 700 channels of Developmental Flight Instrumentation (DFI). Significant insight was gained in the areas of thrust oscillation, vibro-acoustics, predicting jet interactions and slag ejection from solid rocket systems with submerged nozzles. The paper outlines results from the Guidance Navigation & Control (GN&C), Thermal, Vibro-acoustic, Structures, Aero, Aero-Acoustic and Trajectory teams.

Smith, R. Marshall↗
NASA NTRS2003

Aerosciences, Aero-Propulsion and Flight Mechanics Technology Development for NASA's Next Generation Launch Technology Program

The Next Generation Launch Technology (NGLT) program, Vehicle Systems Research and Technology (VSR&T) project is pursuing technology advancements in aerothermodynamics, aeropropulsion and flight mechanics to enable development of future reusable launch vehicle (RLV) systems. The current design trade space includes rocket-propelled, hypersonic airbreathing and hybrid systems in two-stage and single-stage configurations. Aerothermodynamics technologies include experimental and computational databases to evaluate stage separation of two-stage vehicles as well as computational and trajectory simulation tools for this problem. Additionally, advancements in high-fidelity computational tools and measurement techniques are being pursued along with the study of flow physics phenomena, such as boundary-layer transition. Aero-propulsion technology development includes scramjet flowpath development and integration, with a current emphasis on hypervelocity (Mach 10 and above) operation, as well as the study of aero-propulsive interactions and the impact on overall vehicle performance. Flight mechanics technology development is focused on advanced guidance, navigation and control (GN&C) algorithms and adaptive flight control systems for both rocket-propelled and airbreathing vehicles.

Cockrell, Charles E., Jr.↗