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At least 415 records · Page 23

All sources NASA NTRS DOE OSTI

NASA NTRS1992

Component test results from the bearing life improvement program for the Space Shuttle Main Engine oxidizer turbopumps

Interim results from a component test program to improve ball bearing life in the Space Shuttle Main Engine oxygen turbopumps are presented. Two specific bearing applications, using liquid oxygen as the bearing coolant, are addressed. The first, the thrust bearing of the low pressure pump, operates at relatively slow speed with predominantly axial load and little temperature rise in the bulk coolant. Testing has demonstrated a very significant reduction in bearing wear by increasing the bearing internal clearance. Heat generation data was obtained that indicates heavy, intermittent cage-to-ball contact occurs, providing a possible explanation for the observed wear. The second application is the turbine end bearings of the high pressure pump. These bearings operate at high speed and load with the possibility of significant coolant vaporization. Tests on production bearings and bearings having modified internal clearance and curvature yielded scattered but generally poor lives. A dramatic improvement was achieved by coating the standard cage with a thin film of fluorinated ethylene propylene and 15 percent molybdenum disulfide. Very promising results have also been obtained by replacing the standard balls with ones made of silicon nitride, especially in combination with the coated cage.

Keba, John E.↗

NASA NTRS1956

The proper combination of lift loadings for least drag on a supersonic wing

Lagrange's method of undetermined multipliers is applied to the problem of properly combining lift loadings for the least drag at a given lift on supersonic wings. The method shows the interference drag between the optimum loading and any loading at the same lift coefficient to be constant. This is an integral form of the criterion established by Robert T. Jones for optimum loadings. The best combination of four loadings on a delta wing with subsonic leading edges is calculated as a numerical example. The loadings considered have finite pressures everywhere on the plan form. Through the sweepback range the optimum combination of the four nonsingular loadings has about the same drag coefficient as a flat plate with leading-edge thrust.

Grant, Frederick C↗

NASA NTRS2000

Development and Testing of Cooled CMCs for High Thermal Flux Applications

Ceramic Matrix Composites (CMCS) offer the potential for significant weight savings and improved performance for a range of propulsion components utilizing refractory materials. This paper describes the fabrication and testing of functionally graded CMCs produced via a low cost process that represents an order of magnitude cost savings over conventionally fabricated CMCS. Test cylinders were fabricated, characterized and evaluated during exposure to high thermal fluxes of up to 10MW/meters squared at the Laser Hardened Materials Evaluation Laboratory (LHMEL). The bulk density of the CMC tubes was approximately 2.2 grams per cubic centimeters. The performance of cryogenically cooled CMCs was compared with uncooled CMCs against similar thermal loads, and fundamental property data collected for this relatively new breed of CMC. Finally, test thrust cells were fabricated from the functionally graded composite and tested using liquid H2 and O2 propellants at NASA Glen.

Patterson , Mark↗

NASA NTRS1951

Tank Investigation of the Grumman JRF-5 Airplane Equipped with Twin Hydro-Skis TED No. NACA DE 357

A tank investigation has been conducted on a 1/8-size powered dynamic model of the Grumman JRF-5 airplane equipped with twin hydro-skis. The results of tests using two types of skis are presented: one had vertical sides joining the top surface to the chine; the other had the top surface faired to the chine to eliminate the vertical sides. Both configurations had satisfactory longitudinal stability although the model had a slightly greater stable elevator range available when the skis without the vertical sides were attached. Free model tests indicated no instability present when one ski emerged before the other. Considerable excess thrust was available at all speeds with either type of skis. A hump gross load-resistance ratio of 3.37 was obtained with the skis with the vertical sides and 3.53 with the other skis. Landing behavior in smooth water with yaw up to 15deg and roll up to 15deg in opposite directions was satisfactory with either type of skis.

Ramsen, John A.↗

NASA NTRS1976

Calculation of the longitudinal aerodynamic characteristics of wing-flap configurations with externally blown flaps

A theoretical investigation was carried out to extend and improve an existing method for predicting the longitudinal characteristics of wing flap configurations with externally blown flaps (EBF). Two potential flow models were incorporated into the prediction method: a wing and flap lifting-surface model and a turbofan engine wake model. The wing-flap model uses a vortex-lattice approach to represent the wing and flaps. The jet wake model consists of a series of closely spaced vortex rings normal to a centerline which may have vertical and lateral curvature to conform to the local flow field beneath the wing and flaps. Comparisons of measured and predicted pressure distributions, span load distributions on each lifting surface, and total lift and pitching moment coefficients on swept and unswept EBF configurations are included. A wide range of thrust coefficients and flap deflection angles is considered at angles of attack up to the onset of stall. Results indicate that overall lift and pitching-moment coefficients are predicted reasonably well over the entire range. The predicted detailed load distributions are qualitatively correct and show the peaked loads at the jet impingement points, but the widths and heights of the load peaks are not consistently predicted.

Mendenhall, M. R.↗

NASA NTRS1980

Low thrust vehicle concept study

Low thrust chemical (hydrogen-oxygen) propulsion systems configured specifically for low acceleration orbit transfer of large space systems were defined. Results indicate that it is cost effective and least risk to combine the OTV and stowed spacecraft in a single 65 K Shuttle. The study shows that the engine for an optimized low thrust stage (1) does not require very low thrust; (2) 1-3 K thrust range appears optimum; (3) thrust transient is not a concern; (4) throttling probably not worthwhile; and (5) multiple thrusters complicate OTV/LSS design and aggravate LSS loads. Regarding the optimum vehicle for low acceleration missions, the single shuttle launch (LSS and expendable OTV) is most cost effective and least risky. Multiple shuttles increase diameter 20%. The space based radar structure short OTV (which maximizes space available for packaged LSS) favors use of torus tank. Propellant tank pressures/vapor residuals are little affected by engine thrust level or number of burns.

Source record↗

NASA NTRS1971

Development of the Algol III solid rocket motor for SCOUT.

The design and performance of a motor developed for the first stage of the NASA SCOUT-D and E launch vehicles are discussed. The motor delivers a 30% higher total impulse and a 35 to 45% higher payload mass capability than its predecessor, the Algol IIB. The motor is 45 in. in diameter, has a length-to-diameter ratio of 8:1 and delivers an average 100,000-lb thrust for an action time of 72 sec. The motor design features a very high volumetrically loaded internal-burning charge of 17% aluminized polybutadiene propellant, a plasma-welded and heat-treated steel alloy case, and an all-ablative plastic nose liner enclosed in a steel shell. The only significant development problem was the grain design tailoring to account for erosive burning effects which occurred in the high-subsonic-Mach-number port. The tests performed on the motor are described.

Felix, B. R.↗

NASA NTRS1975

The effects on propulsion-induced aerodynamic forces of vectoring a partial-span rectangular jet at Mach numbers from 0.40 to 1.20

An investigation was conducted in the Langley 16-foot transonic tunnel to determine the induced lift characteristics of a vectored thrust concept in which a rectangular jet exhaust nozzle was located in the fuselage at the wing trailing edge. The effects of nozzle deflection angles of 0 deg to 45 deg were studied at Mach numbers from 0.4 to 1.2, at angles of attack up to 14 deg, and with thrust coefficients up to 0.35. Separate force balances were used to determine total aerodynamic and thrust forces as well as thrust forces which allowed a direct measurement of jet turning angle at forward speeds. Wing pressure loading and flow characteristics using oil flow techniques were also studied.

Capone, F. J.↗

NASA NTRS1976

Developments in steady and unsteady aerodynamics for use in aeroelastic analysis and design

A review is given of seven research projects which are aimed at improving the generality, accuracy, and computational efficiency of steady and unsteady aerodynamic theory for use in aeroelastic analysis and design. These projects indicate three major thrusts of current research efforts: (1) more realistic representation of steady and unsteady subsonic and supersonic loads on aircraft configurations of general shape with emphasis on structural-design applications, (2) unsteady aerodynamics for application in active-controls analyses, and (3) unsteady aerodynamics for the frequently critical transonic speed range. The review of each project includes theoretical background, description of capabilities, results of application, current status, and plans for further development and use.

Yates, E. C., Jr.↗

NASA NTRS1981

Structural analysis of cylindrical thrust chambers, volume 3

A system of three computer programs is described for use in conjunction with the BOPAGE finite element program. The programs are demonstrated by analyzing cumulative plastic deformation in a regeneratively cooled rocket thrust chamber. The codes provide the capability to predict geometric and material nonlinear behavior of cyclically loaded structures without performing a cycle-by-cycle analysis over the life of the structure. The program set consists of a BOPACE restart tape reader routine, and extrapolation program and a plot package.

Pearson, M. L.↗

NASA NTRS1984

Electrodynamic Tethers. 1: Power Generator in LEO. 2: Thrust for Propulsion and Power Storage

An electrodynamic tether consists of a long insulated wire in space whose orbital motion cuts across lines of magnetic flux to produce an induce voltage that in typical low orbits averages about 200 v/km. Such a system should be capable of generating substantial electrical power, at the expense of IXB drag acting on its orbital energy. If a reverse current is driven against the induced voltage, the system should act as a motor producing IXB thrust. A reference system was designed, capable of generating 20 KW of power into an electrical load located anywhere along the wire at the expense of 2.6N (20,000 J/sec) drag on the wire. In an ideal system, the conversion between mechanical and electrical energy would reach 100% efficiency. In the actual system part of the 20 KW is lost to internal resistance of the wire, plasma and ionosphere, while the drag force is increased by residual air drag. The 20 KW PMG system as designed is estimated to provide 18.7 KW net power to the load at total drag loss of 20.4 KJ/sec, or an overall efficiency of 92%. Similar systems using heavier wire appear capable of producing power levels in excess of 1 Megawatt at voltages of 2-4 KV, with conversion efficiency between mechanical and electrical power better than 95%. The hollow cathode based system should be readily reversible from generator to motor operation by driving a reverse current using onboard power.

Mccoy, J. E.↗

NASA NTRS1986

Aerodynamic characteristics of several current helicopter tail boom cross sections including the effect of spoilers

Aerodynamic characteristics were determined of three cylindrical shapes representative of tail boom cross sections of the U.S. Army AH-64, UH-60, and UH-1H helicopters. Forces and pressures were measured in a wind-tunnel investigation at the Langley Research Center. Data were obtained for a flow incidence range from -45 to 90 deg and a dynamic pressure range from 1.5 to 50 psf. These ranges provided data representative of full-scale Reynolds numbers and the full range of flow incidence to which these helicopter tail boom shapes would be subjected at low flight speeds. The effects of protuberances such as tail rotor drive-shaft covers and spoilers were evaluated. The data indicate that significant side loads on tail booms of helicopters can be generated and that the addition of spoilers can beneficially alter the side loads. Although an increase in vertical drag occurs, the net effect through reduction of tail rotor thrust required can be an improvement in helicopter performance.

Wilson, J. C.↗

NASA NTRS2002

Fail-Safe Operation of a High-Temperature Magnetic Bearing Investigated for Gas Turbine Engine Applications

The Structural Mechanics and Dynamics Branch at the NASA Glenn Research Center has developed a three-axis high-temperature magnetic bearing suspension rig to enhance the safety of the bearing system up to 1000 F. This test rig can accommodate thrust and radial bearings up to a 22.84 cm (9 in.) diameter with a maximum axial loading of 22.25 kN (5000 lb) and a maximum radial loading up to 4.45 kN (1000 lb). The test facility was set up to test magnetic bearings under high-temperature (1100 F) and high-speed (20,000 rpm) conditions. The magnetic bearing is located at the center of gravity of the rotor between two high-temperature grease-packed mechanical ball bearings. The drive-end duplex angular contact ball bearing, which is in full contact, acts as a moment release and provides axial stability. The outboard end ball bearing has a 0.015-in. radial clearance between the rotor to act as a backup bearing and to compensate for axial thermal expansion. There is a 0.020-in. radial air gap between the stator pole and the rotor. The stator was wrapped with three 1-kW band heaters to create a localized hot section; the mechanical ball bearings were outside this section. Eight threaded rods supported the stator. These incorporated a plunger and Bellville washers to compensate for radial thermal expansion and provide rotor-to-stator alignment. The stator was instrumented with thermocouples and a current sensor for each coil. Eight air-cooled position sensors were mounted outside the hot section to monitor the rotor. Another sensor monitored this rotation of the outboard backup bearing. Ground fault circuit interrupts were incorporated into all power amplifier loops for personnel safety. All instrumentation was monitored and recorded on a LabView-based data acquisition system. Currently, this 12-pole heteropolar magnetic bearing has 13 thermal cycles and over 26 hr of operation at 1000 F.

Choi, Benjamin B.↗

NASA NTRS2010

Stage Separation Failure: Model Based Diagnostics and Prognostics

Safety of the next-generation space flight vehicles requires development of an in-flight Failure Detection and Prognostic (FD&P) system. Development of such system is challenging task that involves analysis of many hard hitting engineering problems across the board. In this paper we report progress in the development of FD&P for the re-contact fault between upper stage nozzle and the inter-stage caused by the first stage and upper stage separation failure. A high-fidelity models and analytical estimations are applied to analyze the following sequence of events: (i) structural dynamics of the nozzle extension during the impact; (ii) structural stability of the deformed nozzle in the presence of the pressure and temperature loads induced by the hot gas flow during engine start up; and (iii) the fault induced thrust changes in the steady burning regime. The diagnostic is based on the measurements of the impact torque. The prognostic is based on the analysis of the correlation between the actuator signal and fault-induced changes in the nozzle structural stability and thrust.

Luchinsky, Dmitry↗

NASA NTRSDate not supplied

Human Mars Mission In-Space Transportation Sensitivity for Nuclear Electric / Chemical Hybrid Propulsion

NASA's Human Exploration and Operation Mission Directorate is continuing to study different concepts and options to field human Mars missions as part of NASA's Moon2Mars directive. In this study, an update to the Hybrid propulsion system was established with the introduction of nuclear electric propulsion system to replace the solar electric propulsion system in an effort to understand the potential impact of a shorter mission to Mars. For crewed missions to Mars, the transportation system sizing is highly dependent on the type of mission and the mission duration. An integrated trajectory analysis capability has been developed and updated to enable this investigation. Longer duration mission may utilize higher efficiency low thrust propulsion system more, but would require higher payload mass due to increase to crew logistics loading and habitation volume. Conversely, shorter duration mission may have lower payload mass, but will require significant increase to propulsion system performance and/or overall system mass. An overall integrated design trade space is defined in this paper to investigate the tradeoff between these scenarios to illuminate the optimality of the transportation system option from a mass perspective. The sensitivity analysis developed in this study will be crucial to understanding the Mars mission design trade space for crewed Mars missions, and will help inform design decisions and investment strategies.

Patrick R. Chai↗

NASA NTRS1974

Cyclic fatigue analysis of rocket thrust chambers. Volume 2: Attitude control thruster high cycle fatigue

A finite element stress analysis was performed for the film cooled throat section of an attitude control thruster. The anlaysis employed the RETSCP finite element computer program. The analysis included thermal and pressure loads, and the effects of temperature dependent material properties, to determine the strain range corresponding to the thruster operating cycle. The configuration and operating conditions considered, correspond to a flightweight integrated thruster assembly which was thrust pulse tested. The computed strain range was used in conjuction with Haynes 188 Universal Slopes minimum life data to predict throat section fatigue life. The computed number of cycles to failure was greater than the number of pulses to which the thruster was experimentally subjected without failure.

Miller, R. W.↗

NASA NTRSDate not supplied

OVERFLOW Analysis of Supersonic Retropropulsion Testing on a Blunt Mars Entry Vehicle Concept

Supersonic retropropulsion (SRP) flow over a Hypersonic Inflatable Aerodynamic Decelerator (HIAD) blunt-body vehicle was simulated using the Overflow Computational Fluid Dynamics (CFD) solver. Two nozzle configurations were tested (1E and 1F) as a subset of the seven total configurations in the Descent System Study (DSS) testing campaign. A generalized, Adaptive Mesh Refinement (AMR) shock capturing and plume refinement technique was developed and calibrated for producing automatic and unique grid systems optimized for any given testing condition. Solution independence from grid resolution was determined by successively increasing grid refinement until asymptotic convergence of the mean aerodynamic loads was observed. Overflow Reynolds-averaged Navier-Stokes (RANS) solutions produced realistic SRP flow phenomena, including Mach disk normal shocks contained in high-thrust plumes and bow shock triple-points. Solutions for the 1E nozzle configuration were steady. Conversely, a subset of the 1F test conditions were unsteady and exhibited periodic, non-sinusoidal expansions and contractions of the streamwise-oriented plume, resulting in unsteady aerodynamic loads on the vehicle. Initial investigations into the effect of turbulence modeling fidelity demonstrated significant differences between the aerodynamic loads simulated with RANS and Detached Eddy Simulation (DES). Multiple flow mechanisms were identified as root causes of these differences, including bow shock shape augmentation and reduced entrainment due to reduced turbulence in the DES simulations. Initial comparisons of pitching moment control authority and aerodynamic drag performance between the tested nozzle configurations demonstrated that the 1E configuration may be more advantageous for the Mars Entry, Descent, and Landing (EDL) task.

Supersonic Retropropulsion↗

NASA NTRSDate not supplied

OVERFLOW Analysis of Supersonic Retropropulsion Testing on a Blunt Mars Entry Vehicle Concept

Supersonic Retropropulsion↗