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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 127 records · Page 7
All sourcesNASA NTRSDOE OSTI
NASA NTRS1973

Development of the propulsion subsystem for the Viking 75 Orbiter

The development of the Viking 75 Orbiter propulsion subsystem has been completed and qualification is under way. This development required an extension of the Mariner Mars 1971 propulsion subsystem technology to meet the Viking 75 requirements of three times the burn time and five times the number of burns. Modifications have been incorporated into the rocket engine to increase the cooling margin for the 2700-second Mars orbit insertion burn. This capability has been demonstrated in both engine firings and subsystem tests. A new surface tension device has been developed to position the propellants (N2O4 and MMH) over the tank outlets during zero-g flight. Extensive scale model testing has been conducted to prove this propellant management concept. This subsystem will provide midcourse corrections, an orbit insertion burn and 20 orbit trims for the Viking 75 Spacecraft. The Viking Project is managed by the Langley Research Center, National Aeronautics and Space Administration.

Vote, F. C.↗
NASA NTRS1980

Conceptual design of an orbital propellant transfer experiment. Volume 2: Study results

The OTV configurations, operations and requirements planned for the period from the 1980's to the 1990's were reviewed and a propellant transfer experiment was designed that would support the needs of these advanced OTV operational concepts. An overall integrated propellant management technology plan for all NASA centers was developed. The preliminary cost estimate (for planning purposes only) is $56.7 M, of which approximately $31.8 M is for shuttle user costs.

Drake, G. L.↗
NASA NTRS1985

GAS payload no. G-025: Study of liquid sloshing behaviour in microgravity

The Get Away Special (GAS) G-025, which flew on shuttle Mission 51-G, examined the behavior of a liquid in a tank under microgravity conditions. The experiment is representative of phenomena occurring in satellite tanks with liquid propellants. A reference fluid in a hemispherical model tank will be subjected to linear acceleration inputs of known levels and frequencies, and the dynamic response of the tank liquid system was recorded. Preliminary analysis of the flight data indicates that the experiment functioned perfectly. The results will validate and refine mathematical models describing the dynamic characteristics of tank-fluid systems. This will in turn support the development of future spacecraft tanks, in particular the design of propellant management devices for surface tension tanks.

C. R. Gilbert↗
NASA NTRS1986

Tethered orbital refueling study

One of the major applications of the space station will be to act as a refueling depot for cryogenic-fueled space-based orbital transfer vehicles (OTV), Earth-storable fueled orbit maneuvering vehicles, and refurbishable satellite spacecraft using hydrazine. One alternative for fuel storage at the space station is a tethered orbital refueling facility (TORF), separated from the space station by a sufficient distance to induce a gravity gradient force that settles the stored fuels. The technical feasibility was examined with the primary focus on the refueling of LO2/LH2 orbital transfer vehicles. Also examined was the tethered facility on the space station. It was compared to a zero-gravity facility. A tethered refueling facility should be considered as a viable alternative to a zero-gravity facility if the zero-gravity fluid transfer technology, such as the propellant management device and no vent fill, proves to be difficult to develop with the required performance.

Fester, Dale A.↗
NASA NTRS1987

Long term cryogenic storage facility systems study

The Long Term Cryogenic Storage Facility Systems Study (LTCSFSS) is a Phase A study of a large capacity propellant depot for the space based, cryogenic orbital transfer vehicle. The study is being performed for Marshall Space Flight Center by General Dynamics Space Systems Division and has five principal objectives: (1) Definition of preliminary concept designs for four storage facility concepts; (2) Selection of preferred concepts through the application of trade studies to candidate propellant management system components; (3) Preparation of a conceptual design for an orbital storage facility; (4) Development of supporting research and technology requirements; and (5) Development of a test program to demonstrate facility performance. The initial study has been completed, and continuation activities are just getting under way to provide greater detail in key areas and accommodate changes in study guidelines and assumptions.

Schuster, John R.↗
NASA NTRS1987

Design and demonstrate the performance of cryogenic components representative of space vehicles: Start basket liquid acquisition device performance analysis

The objective was to design, fabricate and test an integrated cryogenic test article incorporating both fluid and thermal propellant management subsystems. A 2.2 m (87 in) diameter aluminum test tank was outfitted with multilayer insulation, helium purge system, low-conductive tank supports, thermodynamic vent system, liquid acquisition device and immersed outflow pump. Tests and analysis performed on the start basket liquid acquisition device and studies of the liquid retention characteristics of fine mesh screens are discussed.

Source record↗
NASA NTRS1987

Lightweight spacecraft propulsion system selection

Various methods of significantly reducing the mass of planetary spacecraft propulsion systems were addressed. Because of the Space Shuttle Challenger accident, the availability of the Space Transportation System for planetary missions may be limited. Mounting a future planetary mission with lower-energy launch vehicles will require a significant reduction in the spacecraft mass. Smaller launch vehicles for planetary missions were considered. In this study, the assumed payload capability for the planetary missions were considered. In this study, the assumed payload capability for the planetary missions was restricted to the Delta and the Atlas-Centaur launch vehicles. Several propulsion and propulsion-related technologies for spacecraft mass reduction were investigated; ion, arcjet, resistojet, and Magneto-Plasma-Dynamic propulsion and earth-storable high-energy propulsion system thrusters were candidates. The propulsion-related technologies included a start-basket Propellant Management Device, a nitrogen tetroxide diaphragm, and a solid-cryogen cooler. Xenon-ion propulsion allowed the greatest mass reduction.

Palaszewski, B.↗
NASA NTRS1989

Human pathway to the solar system in the 21st century

Some proposed space exploration missions are examined. Consideration is given to requirements for human expeditions to Phobos and Mars and for establishing a lunar observatory and a permanent, self-sufficient facility on Mars. Earth-to-orbit transportation, propellant management, extraterrestrial construction and operations, life support technology, space transfer systems, and automation and robotics are examined in terms of proposed missions.

O'Handley, Douglas A.↗
NASA NTRS1991

Modeling of jet-induced geyser formation in a reduced gravity environment

Flow patterns predicted by a computational model of jet-induced geyser formation in a reduced gravity environment are presented and comparison is made to patterns predicted by experimentally based correlations. The configuration studied is an idealization of a forthcoming flight experiment to examine cryogenic propellant management issues. A transitional version of the ECLIPSE code used as a computational tool for the analyses is described. It is shown that computationally predicted flow patterns are in qualitative agreement with the correlation-based predictions, and some details of the predicted flow fields are given.

Wendl, M. C.↗
NASA NTRS1991

Fuel Systems Architecture (FSA) evaluation criteria and concept evaluation methodology

Consideration is given to two methods developed for the evaluation, screening, and ranking of concepts for Space Exploration Initiative vehicle propellant management systems. The methods selected for handling this multicriteria decision problem are based on the utility theory which transforms both qualitative and quantitative criteria into a nondimensional utility scale for comparison of dissimilar figures of merit. The development of the resultant FSA evaluation criteria and concept evaluation methodology is summarized.

Hendershot, J. E.↗
NASA NTRS1997

Vented Tank Resupply Experiment--Flight Test Results

This paper reports the results of the Vented Tank Resupply Experiment (VTRE) which was flown as a payload on STS 77. VTRE looks at the ability of vane Propellant Management Devices (PMD) to separate liquid and gas in low gravity. VTRE used two clear 0.8 cubic foot tanks one spherical and one with a short barrel section and transferred Refrigerant 113 between them as well as venting it to space. Tests included retention of liquid during transfer, liquid free venting, and recovery of liquid into the PMD after thruster firing. Liquid was retained successfully at the highest flow rate tested (2.73 gpm). Liquid free vents were achieved for both tanks, although at a higher flow rate (0.1591 cfm) for the spherical tank than the other (0.0400 cfm). Recovery from a thruster firing which moved the liquid to the opposite end of the tank from the PMD was achieved in 30 seconds.

David J. Chato↗
NASA NTRS1998

Liquid Motion in a Rotating Tank Experiment (LME)

The Liquid Motion Experiment (LME), which flew on STS 84 in May 1997, was an investigation of liquid motions in spinning, nutating tanks. LME was designed to quantify the effects of such liquid motions on the stability of spinning spacecraft, which are known to be adversely affected by the energy dissipated by the liquid motions. The LME hardware was essentially a spin table which could be forced to nutate at specified frequencies at a constant cone angle, independently of the spin rate. Cylindrical and spherical test tanks, partially filled with liquids of different viscosities, were located at the periphery of the spin table to simulate a spacecraft with off-axis propellant tanks; one set of tanks contained generic propellant management devices (PMDs). The primary quantitative data from the flight tests were the liquid-induced torques exerted on the tanks about radial and tangential axes through the center of the tank. Visual recordings of the liquid oscillations also provided qualitative information. The flight program incorporated two types of tests: sine sweep tests, in which the spin rate was held constant and the nutation frequency varied over a wide range; and sine dwell test, in which both the spin rate and the nutation frequency were held constant. The sine sweep tests were meant to investigate all the prominent liquid resonant oscillations and the damping of the resonances, and the sine dwell tests were meant to quantify the viscous energy dissipation rate of the liquid oscillations for steady state conditions. The LME flight data were compared to analytical results obtained from two companion IR&D programs at Southwest Research Institute. The comparisons indicated that the models predicted the observed liquid resonances, damping, and energy dissipation rates for many test conditions but not for all. It was concluded that improved models and CFD simulations are needed to resolve the differences. This work is ongoing under a current IR&D program.

Deffenbaugh, D. M.↗
NASA NTRS1998

Liquid Motion Experiment Flight Test Results

The Liquid Motion Experiment (LME), designed to study the effects of liquid motion in rotating tanks, was flown on STS 84. LME was essentially a spin table that created a realistic nutation motion of scale-model tanks containing liquid. TWo spherical and two cylindrical transparent tanks were tested simultaneously, and three sets of such tanks were employed to vary liquid viscosity, fill level, and propellant management device (PMD) design. All the tanks were approximately 4.5 inches diameter. The primary test measurements were the radial and tangential torques exerted on the tanks by the liquid. Resonant frequencies and damping of the liquid oscillations were determined by sine sweep tests. For a given tank shape, the resonant frequency depended on fill level. For the cylindrical tanks, the resonances had somewhat different frequencies for the tangential axis (0.55 to 0.75 times spin rate) and the radial axis (0.73 to 0.78 times spin rate), and the tangential axis resonance agreed more closely with available analytical models. For the spherical tanks, the resonant frequencies were between 0.74 to 0.77 times the spin rate and were the same for the tangential and radial axes. The damping coefficients varied from about I% to 3% of critical, depending on tank shape, fill level, and liquid viscosity. 'Me viscous energy dissipation rates of the liquid oscillations were determined from sine dwell tests. The LME energy dissipation rates varied from 0.3 to 0.5 times the estimates obtained from scaling previous ground tests and spacecraft flight data. The PNDs sometimes enhanced the resonances and energy dissipation rates and sometimes decreased them, which points out the need to understand better the effects of PMD on liquid motion as a function of PMD and tank design.

Chato David J.↗
NASA NTRS1999

A Comparison of Two Different Approaches to Hydrazine Loading of Spacecraft: The Use of SCAPE and Alternative Approaches

The loading of spacecraft with Hydrazine type fuels has long been recognized as a hazardous operation. This has led to safety strategies that include the use of SCAPE protective suits for personnel. The use of SCAPE suits have an excellent safety record, however there are associated drawbacks. Drawbacks include the high cost of maintaining and cleaning the suits, reduced mobility and dexterity when wearing the suits, the requirement for extensive specialized health and safety training, and the need to rotate personnel every two hours. A study was undertaken to look at procedures and/or equipment to eliminate or reduce the time spent in SCAPE-type operations. The major conclusions are drawn from observations of the loading of the JPL/NASA spacecraft Deep Space One (DS1) at KSC and the loading of a commercial communications satellite by Motorola at Vandenberg AF Base. The DS1 operations require extensive use of SCAPE suits, while the Motorola operation uses only SPLASH attire with a two-man team on standby in SCAPE. The Motorola team used very different loading equipment and procedures based on an integrated approach involving the propellant supplier. Overall, the Motorola approach was very clean, much faster and simpler than the DS1 procedure. The DS1 spacecraft used a bladder in the propellant tank, whereas the Motorola spacecraft used a Propellant Management Device (PMD). The Motorola approach cannot be used for tanks with bladders. To overcome this problem, some new procedures and new equipment are proposed to enable tanks with bladders to be loaded without using SCAPE, using a modified Motorola approach. Overall, it appears feasible to adopt the non-SCAPE approach while maintaining a very high degree of safety and reliability.

Houseman, John↗
NASA NTRS2001

Trajectory Design and Control for the Compton Gamma Ray Observatory Re-Entry

The Compton Gamma Ray Observatory (CGRO) controlled re-entry operation was successfully conducted in June of 2000. The surviving parts of the spacecraft landed in the Pacific Ocean within the nominal impact target zone. The design of the maneuvers to control the trajectory to accomplish this re-entry presented several challenges. These challenges included the timing and duration of the maneuvers, propellant management, post-maneuver state determination, collision avoidance with other spacecraft, accounting for the break-up of the spacecraft into several pieces with a wide range of ballistic coefficients, and ensuring that the impact footprint would remain within the desired impact target zone in the event of contingencies. This paper presents the initial re-entry trajectory design and traces the evolution of that design into the maneuver sequence used for the re-entry. The paper also discusses the spacecraft systems and operational constraints imposed on the trajectory design and the required modifications to the initial design based on those constraints. Data from the reentry operation are also presented.

Hoge, Susan↗
NASA NTRS1999

Results of Liquid Motion Experiment Analyzed

About half of all scientific and commercial spacecraft spin during some part of their mission. Although spinning has many benefits (increasing stability, controlling the location of liquid propellants, and distributing solar heat loads), it also creates problems because a precession (or wobble) motion is unavoidable. For modern spacecraft, by far the largest source of wobble is energy dissipation in the liquid of partially filled tanks. The liquid s energy dissipation cannot, however, be quantified adequately by any ground test. Current analytical models are also inadequate because fundamental data on fluid motion in low gravity are needed to validate them. Consequently, spacecraft attitude-control systems are designed and operated very conservatively. Nonetheless, spacecraft often still perform poorly in orbit, and some have been lost because of a rapid increase of the wobble rate. The Liquid Motion Experiment (LME) was designed to provide spacecraft designers accurate data on the wobble dynamics of spacecraft that contain large quantities of mobile liquids. LME, which was flown on the space shuttle mission STS-84, was built under contract to the NASA Lewis Research Center by the Southwest Research Institute of San Antonio, Texas. Major accomplishments for 1998 include reduction of the flight data and publication of the experimental results. LME was essentially a spin table that created a realistic nutation motion of scale-model tanks containing liquid. Two spherical and two cylindrical transparent tanks were tested simultaneously, and three sets of such tanks were employed to vary liquid viscosity, fill level, and propellant management device (PMD) design. All the tanks were approximately 4.5 in. in diameter. The primary test measurements were the radial and tangential torques exerted on the tanks by the liquid. These torques could not be measured on the ground because of the masking effects of gravity.

Chato, David J.↗
NASA NTRS2005

Single-String Integration Test Measurements of the NEXT Ion Engine Plume

Measurements were made of a 40 cm ion-thruster plume as part of the single-string-integration-test (SSIT) activity of Phase I of the NASA's Evolutionary Xenon Thruster (NEXT) project. The NEXT ion engine incorporates design improvements that extend NSTAR power levels and efficiencies. During SSIT, an engineering model (EM2) 40 cm engine was operated using an advanced xenon propellant system in combination with either a GRC power console or advanced power processing unit. Integral goals of the single-string phase were to characterize engine performance over the full input power range and to detail thruster operation within the specification of the NEXT throttle table. Plume diagnostics measurements of relative Xe(+) and Xe(++) currents were made using near-field and far-field ExB probes. Planar geometry faraday probes were used to obtain beam current density profiles. This paper reports on the characterization of the EM2 plume over a range of SSIT operating conditions, first with the advanced propellant management system teamed with the GRC power console and then with the power-processing unit.

Snyder, Aaron↗
NASA NTRS2007

A Process for Producing Highly Wettable Aluminum 6061 Surfaces Compatible with Hydrazine

NASA's Global Precipitation Measurement (GPM) mission is an ongoing Goddard Space Flight Center (GSFC) project whose basic objective is to improve global precipitation measurements. The space-based portion of the mission architecture consists of a primary or core spacecraft and a constellation of NASA and contributed spacecrafts. The efforts described in this paper refer to the core spacecraft (hereafter referred to as simply GPM) which is to be fabricated at GSFC. It has been decided that the GPM spacecraft is to be a "design-for-demise-spacecraft." This requirement resulted in the need for a propellant tank that would also demise or ablate to an appropriate degree upon re-entry. Composite overwrapped aluminum lined propellant tanks with aluminum propellant management devices (PMD) were shown by analyses to demise and thus became the baseline configuration for GPM. As part of the GPM tank development effort, long term compatibility and wettability testing with hydrazine was performed on Al6061 and 2219 coupons fabricated and cleaned by conventional processes. Long term compatibility was confirmed. However, the wettability of the aluminum as measured by contact angle produced higher than desired angles (greater than 30 deg.) with excessive scatter. The availability of PMD materials exhibiting consistently low contact angles aids in the design of simple PMDs. Two efforts performed by Angeles Crest Engineering and funded by GSFC were undertaken to reduce the risk of using aluminum for the GPM PMD. The goal of the first effort was to develop a cleaning or treatment process to produce consistently low contact angles. The goal of the second effort was to prove via testing that the processed aluminum would retain compatibility with hydrazine and retain low contact angle after long term exposure to hydrazine. Both goals were achieved. This paper describes both efforts and the results achieved.

Moore, N. R.↗