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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 757 records · Page 42
All sourcesNASA NTRSDOE OSTI
NASA NTRS1985

Direct Transient Analysis of a Fuze Assembly by Axisymmetric Solid Elements

A fuze assembly, which consists of three major parts, nose, collar and sleeve, was designed to survive severe transverse impact giving a maximum base acceleration of 20.000 G. It is shown that hoop failure occurred in the collar after the impact. They also showed that by bonding the collar to the nose, the collar was able to survive the same impact. To find out the effectiveness of the bonding quantitatively, axisymmetric solid elements TRAPAX and TRIAAX were used in modelling the fuze and direct transient analysis was performed. The dynamic stresses in selected elements on the bonded and unbonded collars were compared. The peak hoop stresses in the unbonded collar were found to be up to three times higher than those in the bonded collar. The NASTRAN results explained the observed hoop failure in the unbonded collar. In addition, static and eigenvalue runs were performed as checks on the models prior to the transient runs. The use of the MPCAX cards and the existence and contributors of the calculated first several nearly identical natural frequencies are addressed.

Dai, C. C.↗
NASA NTRS1984

Numerical computation of space shuttle orbiter flow field

The complete inviscid viscous real gas flow around the Space Shuttle Orbiter was computed. Real gas effects are important in predicting the reentry environment around the Orbiter because the high temperatures within the shock layer cause the air to dissociate and ionize, thus invalidating the perfect gas assumption. It is shown that real gas effects has a significant influence on the Orbiter aerodynamics. The approach utilizes a time dependent Navier-Stokes code to compute the subsonic nose portion of the flow field. This nose solution provides the initial conditions for the parabolized Navier-Stokes (PNS) code.

Tannehill, J. C.↗
NASA NTRS1984

A flying superconducting magnet and cryostat for magnetic suspension of wind-tunnel models

The engineering practicality of a persistent high-field superconducting solenoid cryostat as a magnetic suspension and balance system (MSBS) for wind-tunnel testing of aircraft and missile models is examined. The test apparatus is a simple solenoid of filamentary NbTi superconductor with a cupronickel matrix. The apparatus, with a length-to-diameter ratio of 6 to 1 and a radius of 32 mm, used a 0.25 mm wire with a critical current of 27 A in an external field of 6 T. The total heat inleak of 150 mW was achieved. Helium boiloff rates were tested over a range of operating conditions, including pitch attitudes from 10 deg nose down to 90 deg nose up; the rate was estimated as low, but the aerodynamic acceptability of venting gaseous helium has not been determined. It is shown that the effectiveness of the concept increases with increasing scale, and performance in excess of that of conventional ferromagnets is achievable with reduction in size and costs, and with aptness to transonic wind-tunnel testing. Detailed specifications and schematics are included.

Britcher, C.↗
NASA NTRS1984

Investigation of the external flow analysis for density measurements at high altitude

The results of analysis performed on the external flow around the shuttle orbiter nose regions at the Shuttle Upper Atmosphere Mass Spectrometer (SUMS) inlet orifice are presented. The purpose of the analysis is to quantitatively characterize the flow conditions to facilitate SUMS flight data reduction and subsequent determination of orbiter aerodynamic force coefficients in the hypersonic rarefied flow regime. Experimental determination of aerodynamic force coefficients requires accurate simultaneous measurement of forces (or acceleration) and dynamic pressure along with independent knowledge of density and velocity. The SUMS provides independent measurement of dynamic pressure; however, it does so indirectly and requires knowledge of the relationship between measured orifice conditions and the dynamic pressure which can only be determined on the basis of molecule or theory for a winged configuration. Monte Carlo direct simulation computer codes were developed for both the flow field solution at the orifice and for the internal orifice flow. These codes were used to study issues associated with geometric modeling of the orbiter nose geometry and the modeling of intermolecular collisions including rotational energy exchange and a preliminary analysis of vibrational excitation and dissociation effects. Data obtained from preliminary simulation runs are presented.

Bienkowski, G. K.↗
NASA NTRS1986

Self indexing latch system

A self-latching jaw assembly is described which includes a pair of jaws for latching a fitting of an associated structure such as the keel fitting of a space telescope. The jaw assembly automatically locates the fitting received on a positioning pedestal and latches it in its original location on the pedestal without need of precision alignment. The first jaw is actuated and moved to the right whereupon a reciprocating plunger, extended from the nose of the jaw senses the hub of the spool. A jaw position responsive switch is then actuated to terminate the drive motor of the jaw. The astronaut will then actuate the drive motor of the jaw and the jaw will move to the left, whereupon the standoff will engage the face of the jaw in a latching position. In the latching position, the noses of the jaws will be under the flanges of the spool.

John C Gibson↗
NASA NTRS1987

Forebody and missile side forces and the time analogy

A review of the 2-D, impulsively started cylinder literature is given to identify a number of physical parameters that can impact the nondimensional time at which vortex asymmetries begin. These parameters are then related to 3-D testing and explain the lack of data repeatability. The implications of these parameters on 3-D computational modeling is also discussed. An ad hoc form of the time analogy is then used to analyze and organize a slender body data base. An important result of the analysis is that the sectional side force in the nose region is a key to understanding the side force behavior for the configuration as a whole. The nose side force at the first pressure station is shown, for the data analyzed, to uniquely determine the rest of the side force development for laminar flow and to correlate with roll orientation independent of angle of attack and Reynolds number.

Hall, Robert M.↗
NASA NTRS1987

Direct simulation of stagnation streamline flow for hypersonic reentry

The direct simulation Monte Carlo (DSMC) method is applied to simulate one-dimensional flow along the stagnation streamline. The freestream conditions considered are those encountered by the nose region of the Space Shuttle Orbiter during the hypersonic reentry. The range of altitudes (75 to 92 km) considered in the present calculations covers continuum to the less rarefied portion of the transition flow regime. The calculations account for thermal as well as the chemical nonequilibrium effects. The attention is focused on the flow structure along the stagnation streamline for different specified shock locations at a given altitude. The effect of shock location on the stagnation-point gas composition and heat transfer is analyzed for specific altitude; consequently, these results are appropriate for hemispherical bodies with different nose radii. Finally, comparison is made between the present DSMC results and the viscous shock-layer calculations, which show good agreement at the lowest altitude.

Dogra, Virendra K.↗
NASA NTRS1987

High angle of attack hypersonic aerodynamics

A new aerodynamics force model is presented which is based on modified Newtonian theory and empirical correlations. The algebraic model was developed for complete vehicles from take off to orbital speeds and for large angles of attack. Predictions are compared to results for a wind tunnel model at a Mach number of 20, and the full scale Shuttle Orbiter for Mach numbers from 0.25 to 20 for angles of attack from 0 to 50 deg. The maximum shuttle orbiter lift/drag at Mach 10 and 20 is 1.85 at 20-deg angle-of-attack. Aerodynamic force predictions are made for a transatmospheric vehicle, which is a derivative of the Shuttle Orbiter, for Mach numbers from 4 to 27 at angles of attack from 5 to 40 deg. Predicted aerodynamic force data indicate that lift/drag ratios of 5.2 at Mach number 10 and 3.6 at Mach number 26 are obtainable. Changes in force coefficients with changes in: nose angle, sweep angle, and (volume exp 2/3)/planform area are quantified for Mach numbers of 10 and 26. Lift/drag ratios increase with decreasing nose angle and (volume exp 2/3)/planform area and increasing wing sweep angle. Lift/drag ratios are independent of these variables for angles of attack in excess of 20 deg at Mach 10 and 30 deg at Mach 26.

Harloff, Gary J.↗
NASA NTRS1987

A study of the TCAS 2 collision avoidance system mounted on a Boeing 737 aircraft

The purpose of this report is to determine the effects of scattering from major aircraft structures on the TCAS 2 collision avoidance system mounted on a Boeing 737. It is found that the major source of scattering for angles of observation above the horizon is the vertical stabilizer and that its effect may be greatly reduced by mounting the TCAS 2 array close to the nose of the aircraft. In addition, by mounting the array close to the nose, the effects of fuselage blockage on the array patterns at elevation angles below the horizon may be greatly reduced in the forward direction.

Grandchamp, B.↗
NASA NTRS1988

Transonic flow modes of an axisymmetric blunt body

An experimental investigation of transonic flow past a plane-nosed circular cylinder with a plane-nosed circular probe extended coaxially ahead is reported. The possibilities of significant transonic drag reduction and the fluid mechanic phenomena which occur are examined. The probe length and diameter and the approaching flow Mach number are the independent variables. The relations which exist among the probe/cylinder geometry, Mach number, and flow field as revealed by measurements of the drag forces acting on the body are explored.

Koenig, Keith↗
NASA NTRS1988

Bayonet for superfluid helium transfer in space

A prototype superfluid helium bayonet for potential space applications has been developed and evaluated with a low heat leak test apparatus. Measured heat leak of the 13 mm (1/2 inch) bayonet pair is 0.21 W at 1.8 K with an uncertainty of +0.09/-0.05 W. Bayonets are fabricated with thin, electron beam (EB) welded tubes which are EB welded to machined nose and flange pieces. Low heat leak structural integrity is provided by a 0.9 mm thickness of filament wound fiberglass-epoxy. Superfluid creep is restricted by KEL-F nose seals which form vacuum-tight extensions to the bayonet cold end pieces.

Mcintosh, G. E.↗
NASA NTRS1988

Design of a ram accelerator mass launch system

The ram accelerator mass launch system has been proposed to greatly reduce the costs of placing acceleration-insensitive payloads into low earth orbit. The ram accelerator is a chemically propelled, impulsive mass launch system capable of efficiently accelerating relatively large masses from velocities of 0.7 km/sec to 10 km/sec. The principles of propulsion are based on those of a conventional supersonic air-breathing ramjet; however the device operates in a somewhat different manner. The payload carrying vehicle resembles the center-body of the ramjet and accelerates through a stationary tube which acts as the outer cowling. The tube is filled with premixed gaseous fuel and oxidizer mixtures that burn in the vicinity of the vehicle's base, producing a thrust which accelerates the vehicle through the tube. This study examines the requirement for placing a 2000 kg vehicle into a 500 km circular orbit with a minimum amount of on-board rocket propellant for orbital maneuvers. The goal is to achieve a 50 pct payload mass fraction. The proposed design requirements have several self-imposed constraints that define the vehicle and tube configurations. Structural considerations on the vehicle and tube wall dictate an upper acceleration limit of 1000 g's and a tube inside diameter of 1.0 m. In-tube propulsive requirements and vehicle structural constraints result in a vehicle diameter of 0.76 m, a total length of 7.5 m and a nose-cone half angle of 7 degrees. An ablating nose-cone constructed from carbon-carbon composite serves as the thermal protection mechanism for atmospheric transit.

Aarnio, Michael↗
NASA NTRS1989

Low-speed, high-lift aerodynamic characteristics of slender, hypersonic accelerator-type configurations

Two investigations were conducted in the Langley 14 by 22 Foot Subsonic Tunnel to determine the low-speed aerodynamic characteristics of a generic hypersonic accelerator-type configuration. The model was a delta wing configuration incorporating a conical forebody, a simulated wrap-around engine package, and a truncated conical aftbody. Six-component force and moment data were obtained over a range of attack from -4 to 30 degrees and for a sideslip range of + or - 20 degrees. In addition to tests of the basic configuration, component build-up tests were conducted; and the effects of power, forebody nose geometry, canard surfaces, fuselage strakes, and engines on the lower surface alone were also determined. Control power available from deflections of wing flaps and aftbody flaps was also investigated and found to be significantly increased during power-on conditions. Large yawing moments resulted from asymmetric flow fields exhibited by the forebody as revealed by both surface pressure data and flow visualization. Increasing nose bluntness reduced the yawing-moment asymmetry, and the addition of a canard eliminated the yawing-moment asymmetry.

Gatlin, Gregory M.↗
NASA NTRS1987

Independent Orbiter Assessment (IOA): Analysis of the landing/deceleration subsystem

The results of the Independent Orbiter Assessment (IOA) of the Failure Modes and Effects Analysis (FMEA) and Critical Items List (CIL) are presented. The IOA approach features a top-down analysis of the hardware to determine failure modes, criticality, and potential critical items. To preserve independence, this analysis was accomplished without reliance upon the results contained within the NASA FMEA/CIL documentation. This report documents the independent analysis results corresponding to the Orbiter Landing/Deceleration Subsystem hardware. The Landing/Deceleration Subsystem is utilized to allow the Orbiter to perform a safe landing, allowing for landing-gear deploy activities, steering and braking control throughout the landing rollout to wheel-stop, and to allow for ground-handling capability during the ground-processing phase of the flight cycle. Specifically, the Landing/Deceleration hardware consists of the following components: Nose Landing Gear (NLG); Main Landing Gear (MLG); Brake and Antiskid (B and AS) Electrical Power Distribution and Controls (EPD and C); Nose Wheel Steering (NWS); and Hydraulics Actuators. Each level of hardware was evaluated and analyzed for possible failure modes and effects. Criticality was assigned based upon the severity of the effect for each failure mode. Due to the lack of redundancy in the Landing/Deceleration Subsystems there is a high number of critical items.

Compton, J. M.↗
NASA NTRS1989

Cold Flow Determination of the Internal Flow Environment Around the Submerged TVC Nozzle for the Space Shuttle SRM

A series of subscale cold flow tests was performed to quantify the gas flow characteristics at the aft end of the Space Shuttle Solid Rocket Motor. This information was used to support the analyses of the redesigned nozzle/case joint. A portion of the thermal loads at the joint are due to the circumferential velocities and pressure gradients caused primarily by the gimbaling of the submerged nose TVC nozzle. When the nozzle centerline is vectored with respect to the motor centerline, asymmetries are set up in the flow field under the submerged nozzle and immediately adjacent to the nozzle/case joint. Specific program objectives included: determination of the effects of nozzle gimbal angle and propellant geometry on the circumferential flow field; measurement of the static pressure and gas velocities in the vicinity of the nozzle/case joint; use of scaling laws to apply the subscale cold flow data to the full scale SRM; and generation of data for use in validation of 3-D computational fluid dynamic, CFD, models of the SRM flow field. These tests were conducted in the NASA Marshall Space Flight Center Airflow Facility with a 7.5 percent scale model of the aft segment of the SRM. Static and dynamic pressures were measured in the model to quantify the flow field. Oil flow data was also acquired to obtain qualitative visual descriptions of the flow field. Nozzle gimbal angles of 0, 3.5, and 7 deg were used with propellant grain configurations corresponding to motor burn times of 0, 9, 19, and 114 seconds. This experimental program was successful in generating velocity and pressure gradient data for the flow field around the submerged nose nozzle of the Space Shuttle SRM at various burn times and gimbal angles. The nature of the flow field adjacent to the nozzle/case joint was determined with oil droplet streaks, and the velocity and pressure gradients were quantified with pitot probes and wall static pressure measurements. The data was applied to the full scale SRM thru a scaling analysis and the results compared well with the 3-D computational fluid dynamics computer model.

Whitesides, R. H.↗
NASA NTRS1990

Spanwise measurements of vertical components of atmospheric turbulence

Correlation and spectrum magnitude estimates are computed for vertical gust velocity measurements at the nose and wing tips of a NASA B-57B aircraft for six level flight, low speed and low altitude runs and are compared with those of the von Karman atmospheric turbulence model extended for spanwise relationships. The distance between the wing tips was 62.6 ft. Airspeeds ranged from about 330 to 400 ft/sec, heights above the ground ranged from near ground level to about 5250 ft. and gust velocity standard deviations ranged from 4.10 to 8.86 ft/sec. Integral scale lengths, determined by matching measured autocorrelation estimates with those of the model, ranged from 410 to 2050 ft. Digital signals derived from piezoelectric sensors provided continuous pressure and airspeed measurements. Some directional acceleration sensitivity of the sensors was eliminated by sensor orientation, and their performance was spectrally verified for the higher frequencies with supplemental onboard piezoresistive sensors. The model appeared to satisfactorily predict the trends of the measured cross-correlations and cross-spectrum magnitudes, particularly between the nose and wing tips. However, the measured magnitude estimates of the cross-spectra between the wing tips exceeded the predicted levels at the higher frequencies. Causes for the additional power across the wing tips were investigated. Vertical gust velocity components evaluated along and lateral to the flight path implied that the frozen-turbulence-field assumption is a suitable approximation.

Sleeper, Robert K.↗
NASA NTRS1990

Analytical study of the origin and behavior of asymmetric vortices

An hypothesis advanced originally to explain computational observations is supported by theoretical considerations: The asymmetric mean flow observed on bodies of revolution at moderate to high angles of attack is the result of a convective instability of an originally symmetric flow to a time-invariant space-fixed disturbance. Additionally, the time-dependent fluctuations characteristic of the flow at higher angles of attack (up to 90 deg) are the result of an absolute instability of an originally steady flow to a small temporal disturbance of finite duration. Within a common domain, the instability mechanisms may coexist. The experimentally confirmed existence of bistable states, wherein the side-force variation with nose roll angle approaches a square-wave distribution, is attributed to the dominant influence of a pair of trailing vortices from the ogival forebody. Their existence is made possible by the appearance of foci of separation in the skin-friction line pattern beyond a critical angle of attack. The extreme sensitivity of the asymmetric flow orientation to nose geometry, demonstrated experimentally, is attributed to the presence of an indeterminate phase in the family of possible solutions for the three-dimensional wave system.

Tobak, Murray↗
NASA NTRS1990

Equilibrium radiative heating tables for aerobraking in the Martian atmosphere

Studies currently underway for Mars missions often envision the use of aerobraking for orbital capture at Mars. These missions generally involve blunt-nosed vehicles to dissipate the excess energy of the interplanetary transfer. Radiative heating may be of importance in these blunt-body flows because of the highly energetic shock layer around the blunt nose. In addition, the Martian atmosphere contains CO2, whose dissociation products are known to include strong radiators. An inviscid, equilibrium, stagnation point, radiation-coupled flow-field code has been developed for investigating blunt-body atmospheric entry. The method has been compared with ground-based and flight data for air, and reasonable agreement has been found. In the present work, the method was applied to a matrix of conditions in the Martian atmosphere. These conditions encompass most trajectories of interest for Mars exploration spacecraft. The predicted equilibrium radiative heating to the stagnation point of the vehicle is presented.

Hartung, Lin C.↗