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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 235 records · Page 13
All sourcesNASA NTRSDOE OSTI
NASA NTRS2018

A Survey of Methods for Computing Best Estimates of Endoatmospheric and Exoatmospheric Trajectories

Beginning with the mathematical prediction of planetary orbits in the early seventeenth century up through the most recent developments in sensor fusion methods, many techniques have emerged that can be employed on the problem of endo and exoatmospheric trajectory estimation. Although early methods were ad hoc, the twentieth century saw the emergence of many systematic approaches to estimation theory that produced a wealth of useful techniques. The broad genesis of estimation theory has resulted in an equally broad array of mathematical principles, methods and vocabulary. Among the fundamental ideas and methods that are briefly touched on are batch and sequential processing, smoothing, estimation, and prediction, sensor fusion, sensor fusion architectures, data association, Bayesian and non Bayesian filtering, the family of Kalman filters, models of the dynamics of the phases of a rocket's flight, and asynchronous, delayed, and asequent data. Along the way, a few trajectory estimation issues are addressed and much of the vocabulary is defined.

Bernard, William P.↗
NASA NTRS1977

Parametric study of ascent performance of a vertically launched hydrogen-fueled single-stage reusable transport

Ascent performance characteristics were studied for a vertical-take-off, horizontal-landing, single-stage-to-orbit transport vehicle powered by hydrogen fuel rocket with a mixture of fixed- and dual-position nozzles. The analysis was made by systematically varying two sets of trajectory similarity parameters based on the propulsive and aerodynamic characteristics of the vehicle and by calculating a trajectory for each combination of the parameters. The propulsion parameters were the initial thrust-weight ratio, engine combination, and the two expansion ratios of the dual-position rocket nozzles. The aerodynamic parameters were the ratio of reference area to initial weight and the ratio of maximum allowable normal force to initial weight. A first-order analysis was carried out to determine the effect on the performance of including the engine mass penalty. This analysis indicates that the configuration with the lowest initial mass for a given payload requires all dual-position nozzles with initial expansion ratio of 50 and a final expansion ratio of 150.

Rehder, J. J.↗
NASA NTRS2020

Advanced Supersonic Parachute Inflation Research Experiment Preflight Trajectory Modeling and Postflight Reconstruction

The Advanced Supersonic Parachute Inflation Research and Experiments (ASPIRE) was a series of sounding rocket flights aimed at understanding the dynamics of supersonic parachutes that are used for Mars robotic applications. The 2012 Mars Science Laboratory (MSL) had a successful deployment of a supersonic parachute, but based on post-flight analysis of parachute margins, the ASPIRE project was created as a risk-reduction program to improve quantification of these margins and qualify a supersonic parachute for Mars 2020, the follow-on mission to MSL. The first sounding rocket (SR01) flight of ASPIRE occurred near Wallops Island, Virginia on Oct. 4, 2017 and demonstrated the successful deployment and inflation of a MSL build-to-print parachute in flight conditions similar to the 2012 MSL mission. The ASPIRE SR02 and SR03 were successful follow-on flights on Mar. 31, 2018 and Sep. 7, 2018 that demonstrated the new, strengthened supersonic parachute designed for the Mars 2020 project. The SR02 and SR03 parachuteswere targeted to 100% and 140% of the expected flight limit load for Mars 2020 to confirm new margins expected from the strengthened parachute. Prior to all flights, a multi-body flight dynamics simulation was developed to predict the parachute dynamics and was used, in conjunction with other tools, to target Mars-relevant flight conditions. After each flight, the on-board data were used to reconstruct the flight trajectory and to validate the pre-flight dynamics simulation. Post-flight analysis showed that all three tests achieved their targeted conditions and pre-flight modeling bounded the key performance metrics for the parachute. This paper describes the flight mechanics simulation, post-flight reconstruction, and the reconciliation process used to validate the flight models.

Soumyo Dutta↗
NASA NTRS1974

NASA plans for space-processing experiments on sounding rockets

Planning for sounding rocket space processing experiments in the second half of 1975 is reported. Currently operational sounding rocket systems are capable of carrying payloads containing more than 50 kg of experimental apparatus on trajectories including between 5 and 10 min of coasting flight above 100 km altitude. When equipped with stabilization systems to nullify angular velocities in their rotational degrees of freedom, these vehicles can provide an environment in which accelerations are less than 1 mm/sec per 2. The cost of flight operations to implement a rocket experiment program will be of the order of 2000 dollars per kg of discretionary payload, and within the next two years newly operational systems are expected to reduce the cost to about 700 dollars per kg.

Bredt, J. H.↗
NASA NTRS2011

Supersonic Retropropulsion Flight Test Concepts

NASA's Exploration Technology Development and Demonstration Program has proposed plans for a series of three sub-scale flight tests at Earth for supersonic retropropulsion, a candidate decelerator technology for future, high-mass Mars missions. The first flight test in this series is intended to be a proof-of-concept test, demonstrating successful initiation and operation of supersonic retropropulsion at conditions that replicate the relevant physics of the aerodynamic-propulsive interactions expected in flight. Five sub-scale flight test article concepts, each designed for launch on sounding rockets, have been developed in consideration of this proof-of-concept flight test. Commercial, off-the-shelf components are utilized as much as possible in each concept. The design merits of the concepts are compared along with their predicted performance for a baseline trajectory. The results of a packaging study and performance-based trade studies indicate that a sounding rocket is a viable launch platform for this proof-of-concept test of supersonic retropropulsion.

Exploration Technology Development and Demonstrati↗
NASA NTRS1998

1-D Numerical Analysis of RBCC Engine Performance

An RBCC engine combines air breathing and rocket engines into a single engine to increase the specific impulse over an entire flight trajectory. Considerable research pertaining to RBCC propulsion was performed during the 1960's and these engines were revisited recently as a candidate propulsion system for either a single-stage-to-orbit (SSTO) or two-stage-to-orbit (TSTO) launch vehicle. There are a variety of RBCC configurations that had been evaluated and new designs are currently under development. However, the basic configuration of all RBCC systems is built around the ejector scramjet engine originally developed for the hypersonic airplane. In this configuration, a rocket engine plays as an ejector in the air-augmented initial acceleration mode, as a fuel injector in scramjet mode and the rocket in all rocket mode for orbital insertion. Computational fluid dynamics (CFD) is a useful tool for the analysis of complex transport processes in various components in RBCC propulsion systems. The objective of the present research was to develop a transient 1-D numerical model that could be used to predict flow behavior throughout a generic RBCC engine following a flight path.

Han, Samuel S.↗
DOE OSTI2025

Acoustic Rocket Signatures Collected by Smartphones

Rockets generate complex acoustic signatures that can be detected over a thousand kilometers from their source. While many far-field acoustic rocket signatures have been collected and released to the public, very few signatures collected at distances less than 100 km are available. This work presents a curated and annotated dataset of acoustic signatures of 243 rocket launches collected by a network of smartphones stationed at distances between 10 and 70 km from the launch sites, resulting in 1089 individual recordings. Due to the frequency dependence of atmospheric attenuation and the relatively short propagation distances, higher-frequency features not preserved in most publicly available data are observed. The signals are time-aligned to allow for different segments of the signal (ignition, launch, trajectory, chronology) to be more easily examined and compared. Initial analysis of the features of these rocket launch stages is performed, observed features are compared to those found in the existing literature, and comparisons between signals from launches of different rocket types are made. The dataset is annotated and made available to the public to aid future analysis of the characteristics and source mechanisms of rocket acoustics as well as applications such as rocket detection and classification models.

33 ADVANCED PROPULSION SYSTEMS↗
NASA NTRS1971

Optimal N-burn multiorbit injection

Procedure for computing gravity losses over n-burn multiorbit escape trajectories of specified final energy and prediction of optimal burn schedules for nuclear rocket escape maneuvers

Singleton, R. L., Jr.↗
NASA NTRS1972

Rocket ascent G-limited moment-balanced optimization program (RAGMOP)

This document describes the RAGMOP (Rocket Ascent G-limited Momentbalanced Optimization Program) computer program for parametric ascent trajectory optimization. RAGMOP computes optimum polynomial-form attitude control histories, launch azimuth, engine burn-time, and gross liftoff weight for space shuttle type vehicles using a search-accelerated, gradient projection parameter optimization technique. The trajectory model available in RAGMOP includes a rotating oblate earth model, the option of input wind tables, discrete and/or continuous throttling for the purposes of limiting the thrust acceleration and/or the maximum dynamic pressure, limitation of the structural load indicators (the product of dynamic pressure with angle-of-attack and sideslip angle), and a wide selection of intermediate and terminal equality constraints.

Lyons, J. T.↗
NASA NTRS1981

Optimal air-breathing launch vehicle design

A generalized two-point boundary problem methodology, similar to techniques used in deterministic optimal control studies, is applied to the design and flight analysis of a two-stage air-breathing launch vehicle. Simultaneous consideration is given to configuration and trajectory by treating geometry, dynamic discontinuities, and time-dependent flight variables all as controls to be optimized with respect to a single mathematical performance measure. While minimizing fuel consumption, inequality constraints are applied to dynamic pressure and specific force. The optimal system fuel consumption and staging Mach number are found to vary little with changes in the inequality constraints due to substantial geometry and trajectory adjustments. Staging, from an air-breathing first stage to a rocket-powered second stage, consistently occurs near Mach 3.5. The dynamic pressure bound has its most pronounced effects on vehicle geometry, particularly the air-breathing propulsion inlet area, and on the first-stage altitude profile. The specific force has its greatest influence on the second-stage thrust history.

Hattis, P. D.↗
NASA NTRS2013

Ion Propulsion: An Enabling Technology for the Dawn Mission

The Dawn mission, part of NASA's Discovery Program, has as its goal the scientific exploration of the two most massive main-belt asteroids, 4 Vesta, and the dwarf planet 1 Ceres. The Dawn spacecraft was launched from the Cape Canaveral Air Force Station on September 27, 2007 on a Delta-II 7925H-9.5 rocket that placed the 1218-kg spacecraft into an Earth-escape (heliocentic) trajectory with an escape velocity of 11 km/s. On-board the spacecraft is an ion propulsion system (IPS) developed at the Jet Propulsion Laboratory which will provide an additional delta-V of approximately 11 km/s for the heliocentric transfers to each body and for all orbit transfers including orbit capture/escape and transition to the various science orbits. Deterministic thrusting to Vesta began in December 2007 and concluded with orbit capture at Vesta in July 2011. The transfer to Vesta included a Mars gravity assist flyby in February 2009 that provided an additional delta-V of 2.6 km/s and was the only post-launch mission delta-V not provided by IPS. The IPS was used during the 14 months at Vesta for all science orbit transfers and then for Vesta escape. Deterministic thrusting for Ceres began in late August 2012 with a planned arrival date at Ceres in early 2015, whereupon IPS will be used for all science orbit transfers. As of January 2013 the IPS has been operated for approximately 28,000 hours, consumed approximately 280 kg of xenon, and provided a delta-V of approximately 7.5 km/s, the most post-launch delta-V of any spacecraft yet flown. IPS performance characteristics are very close to the expected performance based on analysis and testing performed pre-launch. Use of the IPS together with a medium-priced launch vehicle enabled this high delta-V mission to be performed within the Discovery Program cost cap. This paper provides an overview of the Dawn IPS and its mission operations through departure for Ceres.

Dawn mission overview↗
NASA NTRS1974

Optimal payload ascent trajectories of winged vehicles

A trajectory analysis is presented which shows the advantage of lifting ascent and the penalties associated with increased wing mass due to increases in aerodynamic loading. The vehicle analyzed is made up of an orbiter, an expendable liquid fuel tank, and two solid rocket motors. Results of the tradeoff are illustrated in a graph of the total vehicle mass plotted against the trajectory constraint. Results indicate that even when structural mass penalties are considered lifting ascent provides a payload delivery.

Martin, J. A.↗
NASA NTRS1997

Near-Optimal Re-Entry Trajectories for Reusable Launch Vehicles

A near-optimal guidance law for the descent trajectory for earth orbit re-entry of a fully reusable single-stage-to-orbit pure rocket launch vehicle is derived. A methodology is developed to investigate using both bank angle and altitude as control variables and selecting parameters that maximize various performance functions. The method is based on the energy-state model of the aircraft equations of motion. The major task of this paper is to obtain optimal re-entry trajectories under a variety of performance goals: minimum time, minimum surface temperature, minimum heating, and maximum heading change; four classes of trajectories were investigated: no banking, optimal left turn banking, optimal right turn banking, and optimal bank chattering. The cost function is in general a weighted sum of all performance goals. In particular, the trade-off between minimizing heat load into the vehicle and maximizing cross range distance is investigated. The results show that the optimization methodology can be used to derive a wide variety of near-optimal trajectories.

Chou, H.-C.↗
NASA NTRS2002

A Plasma Rocket Demonstration on the International Space Station

The Advanced Space Propulsion Laboratory at the NASA Johnson Space Center has been engaged in the development of a magneto-plasma rocket for several years. This type of rocket could be used in the future to propel interplanetary spacecraft. One feature of this concept is the ability to vary its specific impulse so that it can be operated in a mode that maximizes propellant efficiency or a mode that maximizes thrust. For this reason the system is called the Variable Specific Impulse Magneto-plasma Rocket or VASIMR. This ability to vary specific impulse and thrust will allow for optimum low thrust interplanetary trajectories and results in shorter trip times than is possible with fixed specific impulse systems while preserving adequate payload margins. In the development of the VASIMR technology, a series of ground-based experiments and space demonstrations are envisioned. A ground-based experiment of a low-power VASIMR prototype rocket is currently underway at the Advanced Space Propulsion Laboratory. The next step is a proposal to build and fly a 25-kilowatt VASIMR rocket as an external payload on the International Space Station. This experiment will provide an opportunity to demonstrate the performance of the rocket in space and measure the induced environment. The experiment will also utilize the space station for its intended purpose as a laboratory with vacuum conditions that cannot be matched by any laboratory on Earth. The VASIMR experiment will also blaze the trail for the wider application of advanced electric propulsion on the space station. An electric propulsion system like VASIMR, if provided with sufficient electrical power, could provide continuous drag force compensation for the space station. Drag compensation would eliminate the need for reboosting the station, an operation that will consume about 60 metric tons of propellant in a ten-year period. In contrast, an electric propulsion system would require very little propellant. In fact, a system like VASIMR can use waste hydrogen from the station's life support system as its propellant. This waste hydrogen is otherwise dumped overboard. Continuous drag compensation would also improve the microgravity conditions on the station. So electric propulsion can reduce propellant delivery requirements and thereby increase available payload capacity and at the same time improve the conditions for scientific research.

Petro, Andrew J.↗
NASA NTRS1972

Parametric analysis of performance and design characteristics for advanced earth-to-orbit shuttles

Performance, trajectory, and design characteristics are presented for (1) a single-stage shuttle with a single advanced rocket engine, (2) a single-stage shuttle with an initial parallel chemical engine and advanced engine burn followed by an advanced engine sustainer burn, (3) a single-stage shuttle with an initial chemical engine burn followed by an advanced engine burn, and (4) a two-stage shuttle with a chemical propulsion booster stage and an advanced propulsion upper stage. The ascent trajectory profile includes a brief initial vertical rise; zero-lift flight through the sensible atmosphere; variational steering into an 83-kilometer by 185-kilometer intermediate orbit; and a fixed, 460-meter per second allowance for subsequent maneuvers. Results are given in terms of burnout mass fractions (including structure and payload), trajectory profiles, propellant loadings, and burn times. These results are generated with a trajectory analysis that includes a parametric variation of the specific impulse from 800 to 3000 seconds and the specific engine weight from 0 to 1.0.

Willis, E. A., Jr.↗
NASA NTRS2010

Tight Fits for Americas Next Moon Rocket, Ares V

America has begun the development of a new heavy lift rocket which will enable humans to return to the moon and reach even farther destinations. Five decades ago, the National Aeronautics and Space Administration designed a system (called Saturn/Apollo) to carry men to the moon and back; the rocket which boosted them to the moon was the Saturn V. Saturn V was huge relative to contemporary rockets and is still the largest rocket ever launched. The new moon rocket is called Ares V. It will insert 40% more payload into low earth orbit than Saturn V; and after docking with the crew spacecraft, it will insert 50% more payload onto the translunar trajectory than Saturn V. The current design of Ares V calls for two liquid-fueled stages and 2 "strap-on" solid rockets. The solid rockets are extended-length versions of the solid rockets used on the Shuttle. The diameter of the liquid stages is at least as large as the first stage of the Saturn V; the height of the lower liquid stage (called the core stage) is longer than the external tank of the Shuttle. Huge rockets require huge infrastructure and, during the Saturn/Apollo era, America invested significantly in manufacturing, assembly and launch facilities which are still in use today. Since the Saturn/Apollo era, America has invested in additional infrastructure for the Shuttle program. Ares V must utilize this existing infrastructure, with reasonable modifications. Building a rocket with 50% more capability in the same buildings, testing it in the same test stands, shipping on the same canals under the same bridges, assembling it in the same building, rolling it to the pad on the same crawler, and launching it from the same launch pad is an engineering and logistics challenge which goes hand-in-hand with designing the structure, tanks, turbines, engines, software, etc. necessary to carry such a large payload to earth orbit and to the moon. This paper quantitatively discusses the significant "tight fits" that are constraining Ares V. The engineers designing and building the infrastructure for the Saturn/Apollo program usually added margins and growth capability; sometimes the size of existing facilities (such as the width of a draw bridge) was not a constraint. Ares V may utilize the "extra" space in the existing facilities and expand other tight fits. Some of the tight fits cannot be overcome without great expense; raising the roof on the Vertical Assembly Building for example. Other tight fits are easily overcome; the transporter at the manufacturing facility for the core stage can pass under low ceilings and later over a dike (without dragging the middle) by retracting or extending the struts which support the stage. Tight fits discussed in this paper include manufacturing (jigs, widths, heights, and local transportation), testing (test stand sizes and crane capability), transportation to the test stands and the launch site (barge, waterway, and rail), assembly (VAB internal dimensions and door size), roll-out limits, and launch pad size.

Jaap, John↗
NASA NTRS2003

An Overview of the NASA Sounding Rocket and Balloon Programs

The U.S. National Aeronautics and Space Administration (NASA) Sounding Rockets and Balloon Programs conduct a total of 50 to 60 missions per year in support of the NASA scientific community. These missions support investigations sponsored by NASA's Offices of Space Science, Life and Microgravity Sciences & Applications, and Earth Science. The Goddard Space Flight Center has management and implementation responsibility for these programs. The NASA Sounding Rockets Program provides the science community with payload development support, environmental testing, launch vehicles, and launch operations from fixed and mobile launch ranges. Sounding rockets continue to provide a cost-effective way to make in situ observations from 50 to 1500 km in the near-earth environment and to uniquely cover the altitude regime between 50 km and 130 km above the Earth's surface. New technology efforts include GPS payload event triggering, tailored trajectories, new vehicle configuration development to expand current capabilities, and the feasibility assessment of an ultra high altitude sounding rocket vehicle. The NASA Balloon Program continues to make advancements and developments in its capabilities for support of the scientific ballooning community. The Long Duration Balloon (LDB) is capable of providing flight durations in excess of two weeks and has had many successful flights since its development. The NASA Balloon Program is currently engaged in the development of the Ultra Long Duration Balloon (ULDB), which will be capable of providing flight times up to 100-days. Additional development efforts are focusing on ultra high altitude balloons, station keeping techniques and planetary balloon technologies.

Eberspeaker, Philip J.↗