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At least 955 records · Page 53
All sourcesNASA NTRSDOE OSTI
NASA NTRS1999

An Air-Breathing Launch Vehicle Concept for Single-Stage-to-Orbit

The "Trailblazer" is a 300-lb payload, single-stage-to-orbit launch vehicle concept that uses air-breathing propulsion to reduce the required propellant fraction. The integration of air-breathing propulsion is done considering performance, structural and volumetric efficiency, complexity, and design risk. The resulting configuration is intended to be viable using near-term materials and structures. The aeropropulsion performance goal for the Trailblazer launch vehicle is an equivalent effective specific impulse (I*) of 500 sec. Preliminary analysis shows that this requires flight in the atmosphere to about Mach 10, and that the gross lift-off weight is 130,000 lb. The Trailblazer configuration and proposed propulsion system operating modes are described. Preliminary performance results are presented, and key technical issues are highlighted. An overview of the proposed program plan is given.

Trefny, Charles J.↗
NASA NTRS1999

High-Area-Ratio Rocket Nozzle at High Combustion Chamber Pressure: Experimental and Analytical Validation

Experimental data were obtained on an optimally contoured nozzle with an area ratio of 1025:1 and on a truncated version of this nozzle with an area ratio of 440:1. The nozzles were tested with gaseous hydrogen and liquid oxygen propellants at combustion chamber pressures of 1800 to 2400 psia and mixture ratios of 3.89 to 6.15. This report compares the experimental performance, heat transfer, and boundary layer total pressure measurements with theoretical predictions of the current Joint Army, Navy, NASA, Air Force (JANNAF) developed methodology. This methodology makes use of the Two-Dimensional Kinetics (TDK) nozzle performance code. Comparisons of the TDK-predicted performance to experimentally attained thrust performance indicated that both the vacuum thrust coefficient and the vacuum specific impulse values were approximately 2.0-percent higher than the turbulent prediction for the 1025:1 configurations, and approximately 0.25-percent higher than the turbulent prediction for the 440:1 configuration. Nozzle wall temperatures were measured on the outside of a thin-walled heat sink nozzle during the test fittings. Nozzle heat fluxes were calculated front the time histories of these temperatures and compared with predictions made with the TDK code. The heat flux values were overpredicted for all cases. The results range from nearly 100 percent at an area ratio of 50 to only approximately 3 percent at an area ratio of 975. Values of the integral of the heat flux as a function of nozzle surface area were also calculated. Comparisons of the experiment with analyses of the heat flux and the heat rate per axial length also show that the experimental values were lower than the predicted value. Three boundary layer rakes mounted on the nozzle exit were used for boundary layer measurements. This arrangement allowed total pressure measurements to be obtained at 14 different distances from the nozzle wall. A comparison of boundary layer total pressure profiles and analytical predictions show good agreement for the first 0.5 in. from the nozzle wall; but the further into the core flow that measurements were taken, the more that TDK overpredicted the boundary layer thickness.

Jankovsky, Robert S.↗
NASA NTRS1999

Non-Toxic Orbital Maneuvering System Engine Development

Recent results using the Aestus engine operated with LOx/ethanol propellant are presented. An experimental program at Rocketdyne Propulsion and Power is underway to adapt this engine for the Boeing Reusable Space Systems Division non-toxic Orbital Maneuvering System/Reaction control System (OMS/RCS) system. Daimler-Chrysler Aerospace designed the Aestus as an nitrogen tetroxide/monomethyl hydrazine (NTO/MMH) upper-stage engine for the Ariane 5. The non-toxic OMS/RCS system's preliminary design requires a LOx/ethanol (O2/C2H5OH) engine that operates with a mixture ratio of 1.8, a specific impulse of 323 seconds, and fits within the original OMS design envelope. This paper describes current efforts to meet these requirements including, investigating engine performance using LOx/ethanol, developing the en-ine system sizing package, and meeting the vehicle operation parameters. Data from hot-fire testing are also presented and discussed.

Green, Christopher↗
NASA NTRS1998

Euler Calculations at Off-Design Conditions for an Inlet of Inward Turning RBCC-SSTO Vehicle

The inviscid performance of an inward turning inlet design is calculated computationally for the first time. Hypersonic vehicle designs based on the inward turning inlets have been shown analytically to have increased effective specific impulse and lower heat load than comparably designed vehicles with two-dimensional inlets. The inward turning inlets are designed inversely from inviscid stream surfaces of known flow fields. The computational study is performed on a Mach 12 inlet design to validate the performance predicted by the design code (HAVDAC) and calculate its off-design Mach number performance. The three-dimensional Euler equations are solved for Mach 4, 8, and 12 using a software package called SAM, which consists of an unstructured mesh generator (SAMmesh), a three-dimensional unstructured mesh flow solver (SAMcfd), and a CAD-based software (SAMcad). The computed momentum averaged inlet throat pressure is within 6% of the design inlet throat pressure. The mass-flux at the inlet throat is also within 7 % of the value predicted by the design code thereby validating the accuracy of the design code. The off-design Mach number results show that flow spillage is minimal, and the variation in the mass capture ratio with Mach number is comparable to an ideal 2-D inlet. The results from the inviscid flow calculations of a Mach 12 inward turning inlet indicate that the inlet design has very good on and off-design performance which makes it a promising design candidate for future air-breathing hypersonic vehicles.

Takashima, N.↗
NASA NTRS1999

Antimatter Production for Near-Term Propulsion Applications

This presentation discusses the use and potential of power generated from Proton-Antiproton Annihilation. The problem is that there is not enough production of anti-protons, and that the production methods are inefficient. The cost for 1 gram of antiprotons is estimated at 62.5 trillion dollars. Applications which require large quantities (i.e., about 1 kg) will require dramatic improvements in the efficiency of the production of the antiprotons. However, applications which involve small quantities (i.e., 1 to 10 micrograms may be practical with a relative expansion of capacities. There are four "conventional" antimatter propulsion concepts which are: (1) the solid core, (2) the gas core, (3) the plasma core, and the (4) beam core. These are compared in terms of specific impulse, propulsive energy utilization and vehicle structure/propellant mass ratio. Antimatter-catalyzed fusion propulsion is also evaluated. The improvements outlined in the presentation to the Fermilab production, and other sites. capability would result in worldwide capacity of several micrograms per year, by the middle of the next decade. The conclusions drawn are: (1) the Conventional antimatter propulsion IS not practical due to large p-bar requirement; (2) Antimatter-catalyzed systems can be reasonably considered this "solves" energy cost problem by employing substantially smaller quantities; (3) With current infrastructure, cost for 1 microgram of p-bars is $62.5 million, but with near-term improvements cost should drop; (4) Milligram-scale facility would require a $15 billion investment, but could produce 1 mg, at $0.1/kW-hr, for $6.25 million.

Gerrish, Harold P.↗
NASA NTRS1999

Multi-Axis Thrust Measurements of the EO-1 Pulsed Plasma Thruster

Pulsed plasma thrusters are low thrust propulsive devices which have a high specific impulse at low power. A pulsed plasma thruster is currently scheduled to fly as an experiment on NASA's Earth Observing-1 satellite mission. The pulsed plasma thruster will be used to replace one of the reaction wheels. As part of the qualification testing of the thruster it is necessary to determine the nominal thrust as a function of charge energy. These data will be used to determine control algorithms. Testing was first completed on a breadboard pulsed plasma thruster to determine nominal or primary axis thrust and associated propellant mass consumption as a function of energy and then later to determine if any significant off-axis thrust component existed. On conclusion that there was a significant off-axis thrust component with the bread-board in the direction of the anode electrode, the test matrix was expanded on the flight hardware to include thrust measurements along all three orthogonal axes. Similar off-axis components were found with the flight unit.

Arrington, Lynn A.↗
NASA NTRS1999

CVD Rhenium Engines for Solar-Thermal Propulsion Systems

Solar-thermal upper-stage propulsion systems have the potential to provide specific impulse approaching 900 seconds, with 760 seconds already demonstrated in ground testing. Such performance levels offer a 100% increase in payload capability compared to state-of-the-art chemical upper-stage systems, at lower cost. Although alternatives such as electric propulsion offer even greater performance, the 6- to 18- month orbital transfer time is a far greater deviation from the state of the art than the one to two months required for solar propulsion. Rhenium metal is the only material that is capable of withstanding the predicted thermal, mechanical, and chemical environment of a solar-thermal propulsion device. Chemical vapor deposition (CVD) is the most well-established and cost-effective process for the fabrication of complex rhenium structures. CVD rhenium engines have been successfully constructed for the Air Force ISUS program (bimodal thrust/electricity) and the NASA Shooting Star program (thrust only), as well as under an Air Force SBIR project (thrust only). The bimodal engine represents a more long-term and versatile approach to solar-thermal propulsion, while the thrust-only engines provide a potentially lower weight/lower cost and more near-term replacement for current upper-stage propulsion systems.

Williams, Brian E.↗
NASA NTRS1999

Rocket-in-a-Duct Performance Analysis

An axisymmetric, 110 N class, rocket configured with a free expansion between the rocket nozzle and a surrounding duct was tested in an altitude simulation facility. The propellants were gaseous hydrogen and gaseous oxygen and the hardware consisted of a heat sink type copper rocket firing through copper ducts of various diameters and lengths. A secondary flow of nitrogen was introduced at the blind end of the duct to mix with the primary rocket mass flow in the duct. This flow was in the range of 0 to 10% of the primary massflow and its effect on nozzle performance was measured. The random measurement errors on thrust and massflow were within +/-1%. One dimensional equilibrium calculations were used to establish the possible theoretical performance of these rocket-in-a-duct nozzles. Although the scale of these tests was small, they simulated the relevant flow expansion physics at a modest experimental cost. Test results indicated that lower performance was obtained at higher free expansion area ratios and longer ducts, while, higher performance was obtained with the addition of secondary flow. There was a discernable peak in specific impulse efficiency at 4% secondary flow. The small scale of these tests resulted in low performance efficiencies, but prior numerical modeling of larger rocket-in-a-duct engines predicted performance that was comparable to that of optimized rocket nozzles. This remains to be proven in large-scale, rocket-in-a-duct tests.

Schneider, Steven J.↗
NASA NTRS1999

External Pulsed Plasma Propulsion and Its Potential for the Near Future

This paper examines External Pulsed Plasma Propulsion (EPPP), a propulsion concept that derives its thrust from plasma waves generated from a series of small, supercritical fission/fusion pulses behind an object in space. For spacecraft applications, a momentum transfer mechanism translates the intense plasma wave energy into a vehicle acceleration that is tolerable to the rest of the spacecraft and its crew. This propulsion concept offers extremely high performance in terms of both specific impulse (Isp) and thrust-to-weight ratio, something that other concepts based on available technology cannot do, The political concerns that suspended work on this type of system (i.e. termination of Project ORION) may now not be as insurmountable as they were in 1965. The appeal of EPPP stems from its relatively low cost and reusability, fast interplanetary transit times, safety and reliability, and independence from major technological breakthroughs. In fact, a first generation EPPP system based on modern-day technology (i.e., GABRIEL - an evolutionary framework- of EPPP concepts) may very well be the only form of propulsion that could realistically be developed to perform ambitious human exploration beyond Mars in the 21st century. It could also provide the most effective approach for deterrence against collision between earth and small planetary objects - a growing concern over recent years.

Bonometti, J. A.↗
NASA NTRS1999

Performance Evaluation of a 50kW Hall Thruster

An experimental investigation was conducted on a laboratory model Hall thruster designed to operate at power levels up to 50 kW. During this investigation the engine's performance was characterized over a range of discharge currents from 10 to 36 A and a range of discharge voltages from 200 to 800 V Operating on the Russian cathode a maximum thrust of 966 mN was measured at 35.6 A and 713.0 V. This corresponded to a specific impulse of 3325 s and an efficiency of 62%. The maximum power the engine was operated at was 25 kW. Additional testing was conducted using a NASA cathode designed for higher current operation. During this testing, thrust over 1 N was measured at 40.2 A and 548.9 V. Several issues related to operation of Hall thrusters at these high powers were encountered.

Jacobson, David T.↗
NASA NTRS1999

Test Results of a 200 W Class Hall Thruster

The performance of a 200 W class Hall thruster was evaluated. Performance measurements were taken at power levels between 90 W and 250 W. At the nominal 200 W design point, the measured thrust was 11.3 mN. and the specific impulse was 1170 s excluding cathode flow in the calculation. A laboratory model 3 mm diameter hollow cathode was used for all testing. The engine was operated on laboratory power supplies in addition to a breadboard power processing unit fabricated from commercially available DC to DC converters.

Jacobson, David↗
NASA NTRS1999

High Pressure Earth Storable Rocket Technology Program-Hipes Options 1/2 Report

Under the High Pressure Earth Storable Rocket Technology (HIPES) Program, TRW successfully completed testing of two 100 lbf thrust class rhenium chambers using N204-MMH. The first chamber was successfully fired for 4789 seconds of operating time with a maximum duration of 700 seconds. This chamber had been previously fired for 5230 seconds with N2O4-N2H4. The second chamber was successfully fired for 8085 seconds with a maximum firing duration of 1200 seconds. The Isp (specific impulse) for both chambers ranged from 323 lbf-sec/lbm to 330 lbf-sec/lbm.

Chazen, M. L.↗
NASA NTRS2000

Atomic-Based-Combined-Cycle Analysis

Atomic-based-combined-cycle (ABCC) engine combines an air-breathing ramjet engine with an atomic reactor to increase the mission-averaged specific impulse and thereby increasing the dry-mass ratio. ABCC engine is similar to RBCC engine except that energy needed for the propulsive power is derived from nuclear reaction rather than chemical combustion used in the RBCC engine. The potential performance improvement of an ABCC engine over a RBCC engine comes from two factors. Firstly, the energy density of nuclear reaction is several order of magnitudes higher than the chemical combustion. Secondly, hydrogen can produce much higher nozzle exit velocity because of its small molecular weight. A one-dimensional, transient numerical model was used to analyze a generic scramjet engine and it is used as a baseline to evaluate an imaginary ABCC engine performance. A nuclear reactor is treated as a black box energy source that replaces the role of the primary rocket and the chemical combustion chamber in a RBCC engine. Hydrogen is heated by the reactor and accelerated to produce high-speed ejection velocity. The ejection velocity up 10,000 m/sec is theoretically possible because of high energy density from the reactor and large gas constant of the hydrogen. Oxygen contained in the entrained air reacts with hydrogen and produces propulsive power for ejector mode operation. To provide enough thrust for initial acceleration, relatively large amount of hydrogen must be pumped through the reactor. Amount of oxygen contained in the entrained air may not be sufficient to burn all hydrogen and consequently combustion could occur at the end of exit nozzle. It is assumed that this combustion process is constant-pressure combustion at 1.0 atmospheric pressure and thus not affects the nozzle exit condition.

Han, Sam↗
NASA NTRS2000

Magnetic Flux Compression Concept for Nuclear Pulse Propulsion and Power

The desire for fast, efficient interplanetary transport requires propulsion systems having short acceleration times and very high specific impulse attributes. Unfortunately, most highly efficient propulsion systems which are within the capabilities of present day technologies are either very heavy or yield very low impulse such that the acceleration time to final velocity is too long to be of lasting interest, One exception, the nuclear thermal thruster, could achieve the desired acceleration but it would require inordinately large mass ratios to reach the range of desired final velocities. An alternative approach, among several competing concepts that are beyond our modern technical capabilities, is a pulsed thermonuclear device utilizing microfusion detonations. In this paper, we examine the feasibility of an innovative magnetic flux compression concept for utilizing microfusion detonations, assuming that such low yield nuclear bursts can be realized in practice. In this concept, a magnetic field is compressed between an expanding detonation driven diamagnetic plasma and a stationary structure formed from a high temperature superconductor (HTSC). In general, we are interested in accomplishing two important functions: (1) collimation of a hot diamagnetic plasma for direct thrust production; and (2) pulse power generation for dense plasma ignition. For the purposes of this research, it is assumed that rnicrofusion detonation technology may become available within a few decades, and that this approach could capitalize on recent advances in inertial confinement fusion ICF) technologies including magnetized target concepts and antimatter initiated nuclear detonations. The charged particle expansion velocity in these detonations can be on the order of 10 (exp 6)- 10 (exp 7) meters per second, and, if effectively collimated by a magnetic nozzle, can yield the Isp and the acceleration levels needed for practical interplanetary spaceflight. The ability to ignite pure fusion micro-bursts with reasonable levels of input energy is an equally challenging scientific problem. It remains to be seen, however, whether an effective ignition driver can be developed which meets the requirements for practical spaceflight application (namely high power density, compactness, low weight, and low cost). In this paper, system level performance and design issues are examined including generator performance, magnetic flux compression processes, magnetic diffusion processes, high temperature superconductor (HTSC) material properties, plasmadynamic processes, detonation plasma expansion processes, magnetohydrodynamic instabilities, magnetic nozzle performance, and thrust production performance. Representative generator performance calculations based on a simplified skin layer formulation are presented as well as the results of exploratory small-scale laboratory experiments on magnetic flux diffusion in HTSC materials. In addition, planned follow-on scientific feasibility experiments are described which utilize high explosive detonations and high energy gas discharges to simulate the plasma conditions associated with thermonuclear micro-detonations.

Litchford, Ronald J.↗
NASA NTRS1999

Overview on NASA's Advanced Electric Propulsion Concepts Activities

Advanced electric propulsion research activities are currently underway that seek to addresses feasibility issues of a wide range of advanced concepts, and may result in the development of technologies that will enable exciting new missions within our solar system and beyond. Each research activity is described in terms of the present focus and potential future applications. Topics include micro-electric thrusters, electrodynamic tethers, high power plasma thrusters and related applications in materials processing, variable specific impulse plasma thrusters, pulsed inductive thrusters, computational techniques for thruster modeling, and advanced electric propulsion missions and systems studies.

Frisbee, Robert H.↗
NASA NTRS2000

Internal Flow Analysis of Large L/D Solid Rocket Motors

Traditionally, Solid Rocket Motor (SRM) internal ballistic performance has been analyzed and predicted with either zero-dimensional (volume filling) codes or one-dimensional ballistics codes. One dimensional simulation of SRM performance is only necessary for ignition modeling, or for motors that have large length to port diameter ratios which exhibit an axial "pressure drop" during the early burn times. This type of prediction works quite well for many types of motors, however, when motor aspect ratios get large, and port to throat ratios get closer to one, two dimensional effects can become significant. The initial propellant grain configuration for the Space Shuttle Reusable Solid Rocket Motor (RSRM) was analyzed with 2-D, steady, axi-symmetric computational fluid dynamics (CFD). The results of the CFD analysis show that the steady-state performance prediction at the initial burn geometry, in general, agrees well with 1-D transient prediction results at an early time, however, significant features of the 2-D flow are captured with the CFD results that would otherwise go unnoticed. Capturing these subtle differences gives a greater confidence to modeling accuracy, and additional insight with which to model secondary internal flow effects like erosive burning. Detailed analysis of the 2-D flowfield has led to the discovery of its hidden 1-D isentropic behavior, and provided the means for a thorough and simplified understanding of internal solid rocket motor flow. Performance parameters such as nozzle stagnation pressure, static pressure drop, characteristic velocity, thrust and specific impulse are discussed in detail and compared for different modeling and prediction methods. The predicted performance using both the 1-D codes and the CFD results are compared with measured data obtained from static tests of the RSRM. The differences and limitations of predictions using ID and 2-D flow fields are discussed and some suggestions for the design of large L/D motors and more critically, motors with port to throat ratios near one, are covered.

Laubacher, Brian A.↗
NASA NTRS2000

Upper Stage Flight Experiment 10K Engine Design and Test Results

A 10,000 lbf thrust chamber was developed for the Upper Stage Flight Experiment (USFE). This thrust chamber uses hydrogen peroxide/JP-8 oxidizer/fuel combination. The thrust chamber comprises an oxidizer dome and manifold, catalyst bed assembly, fuel injector, and chamber/nozzle assembly. Testing of the engine was done at NASA's Stennis Space Center (SSC) to verify its performance and life for future upper stage or Reusable Launch Vehicle applications. Various combinations of silver screen catalyst beds, fuel injectors, and combustion chambers were tested. Results of the tests showed high C* efficiencies (97% - 100%) and vacuum specific impulses of 275 - 298 seconds. With fuel film cooling, heating rates were low enough that the silica/quartz phenolic throat experienced minimal erosion. Mission derived requirements were met, along with a perfect safety record.

Ross, R.↗
NASA NTRS1999

Recent Progress on the VASIMR

Experimental and theoretical studies on the Variable Specific Impulse Magnetoplasma Rocket (VASIMR) have continued through a NASA led collaborative program involving several research groups. In the experimental area, performance characterization of the VASIMR helicon plasma source has been obtained over a portion of the parameter space, with helium and hydrogen propellant. Density (10(exp 18) - 10(exp 19)/ cubic meter) and temperature (5 eV) were measured at moderate degree of ionization in two separate experimental devices. Helicon design improvement and optimization will be discussed. Experiments with the ion cyclotron resonance heating (ICRH) subsection have begun and preliminary results will be discussed. Theoretical picture and integrated numerical simulation continue to be refined to account for the main physics elements of the VASIMR, including RF absorption and particle acceleration with subsequent detachment in the magnetic nozzle.

ChangDiaz, Franklin R.↗