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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 163 records · Page 9
All sourcesNASA NTRSDOE OSTI
NASA NTRS2016

What Would Happen to Superstorm Sandy Under the Influence of a Substantially Warmer Atlantic Ocean?

Based on ensemble numerical simulations, we find that possible responses of Sandy-like superstorms under the influence of a substantially warmer Atlantic Ocean bifurcate into two groups. In the first group, storms are similar to present-day Sandy from genesis to extratropical transition, except they are much stronger, with peak Power Destructive Index (PDI) increased by 50-80%, heavy rain by 30-50%, and maximum storm size (MSS) approximately doubled. In the second group, storms amplify substantially over the interior of the Atlantic warm pool, with peak PDI increased by 100-160%, heavy rain by 70-180%, and MSS more than tripled compared to present-day Superstorm Sandy. These storms when exiting the warm pool, recurve northeastward out to sea, subsequently interact with the developing midlatitude storm by mutual counterclockwise rotation around each other and eventually amplify into a severe Northeastern coastal storm, making landfall over the extreme northeastern regions from Maine to Nova Scotia.

radar reflectivities↗
NASA NTRS2004

Procedure for Determining One-Dimensional Flow Distributions in Arbitrarily Connected Passages Without the Influence of Pumping

A calculation procedure is presented which allows the one-dimensional determination of flow distributions in arbitrarily connected (branching) flow passages having multiple inlets and exits. The procedure uses an adaptation of the finite element technique, iteratively coupled with an accurate one-dimensional flow solver. The procedure eliminates the usual restrictions inherent with finite element flow calculations. Unlike existing one-dimensional methods, which require simplifications to the flow equations (uncoupling the momentum and energy equations), to allow for arbitrary branching and multiple inlets and exits, the only limitation of the described methodology is that, at present, it can only accommodate non-rotating configurations (no pumping effects). The calculation procedure is robust, and will always converge for physically possible flow. The procedure is described, and its use is illustrated by an example.

Meitner, Peter L.↗
NASA NTRS1959

Effect of Multiple-Jets Exits on the Base Pressure of a Simple Wing-Body Combination at Mach Numbers of 0.6 to 1.27

An investigation has been conducted at Mach numbers of 0.6 to 1.27 to determine the effect of multiple-jet exits on the base pressure of a simple wing-body combination. The design Mach number of the nozzles ranged from 1 to 3 at jet exit diameters equal to 36.4 to 75 percent of the model thickness. Jet total-pressure to free-stream static-pressure ratios ranged from 1 (no flow) to 34.2. The results show that the variation of base pressure coefficient with jet pressure ratio for the model tested was similar to that obtained for single nozzles in bodies of revolution in other investigations. As in the case for single jets the base pressure coefficient for the present model became less negative as the jet exit diameter increased. For a constant throat diameter and an assumed schedule of jet pressure ratio over the speed range of these tests, nozzle Mach number had only a small effect on base pressure coefficient.

Cubbage, James M., Jr.↗
NASA NTRS1986

SRB/SLEEC (Solid Rocket Booster/Shingle Lap Extendible Exit Cone) feasibility study, volume 2. Appendix A: Design study for a SLEEC actuation system

The results are presented of a design feasibility study of a self-contained (powered) actuation system for a Shingle Lap Extendible Exit Cone (SLEEC) for Transportation System (STS). The evolution of the SLEEC actuation system design is reviewed, the final design concept is summarized, and the results of the detailed study of the final concept of the actuation system are treated. A conservative design using proven mechanical components was established as a major program priority. The final mechanical design has a very low development risk since the components, which consist of ballscrews, gearing, flexible shaft drives, and aircraft cables, have extensive aerospace applications and a history of proven reliability. The mathematical model studies have shown that little or no power is required to deploy the SLEEC actuation system because acceleration forces and internal pressure from the rocket plume provide the required energies. A speed control brake is incorporated in the design in order to control the rate of deployment.

Thompson, D. S.↗
NASA NTRS1993

Streamwise vorticity generation and mixing enhancement in free jets by 'delta-tabs'

The effect of triangular tabs, placed at the nozzle exit, on the evolution of free jets is investigated. The effect, a large distortion of the jet cross section and a resultant increase in mixing downstream, has been inferred before to be due to a pair of streamwise vortices originating from each tab. In this paper, the generation mechanism of the streamwise vorticity (omega sub x) is considered first. Two sources are postulated. One is the upstream 'pressure hill', produced by the tab, which appears to be the dominant source. Another is due to vortex filaments shed from the sides of the tab and reoriented downstream by the mean shear of the mixing layer. In the case of a 'delta-tab', a triangular tab with its apex leaning downstream, vorticity from the two sources explain the stronger effect in that configuration. Data on the vorticity evolution for the effect of two delta-tabs are presented, up to twelve jet diameters from the exit, which show that the streamwise vortices persist even at the farthest measurement station. The magnitude of omega sub x-maximum decays continually with distance from the nozzle, its ratio to azimuthal vorticity maximum is found to be about 1/5 everywhere. The relative effect of a delta-tab on jets from an axisymmetric nozzle and a 8:1 rectangular nozzle is also studied. The mixing layer distortion is found to be less pronounced in the rectangular case. The jet mixing, as manifested by the mass flux measured at a downstream station, is increased in the axisymmetric jet but it is decreased in the rectangular jet under consideration by the delta-tab.

Zaman, K. B. M. Q.↗
NASA NTRS1993

Streamwise vorticity generation and mixing enhancement in free jets by delta-tabs

The effect of triangular tabs, placed at the nozzle exit, on the evolution of free jets is investigated. The effect, a large distortion of the jet cross section and a resultant increase in mixing downstream, has been inferred before to be due to a pair of streamwise vortices originating from each tab. In this paper, the generation mechanism of the stream wise vorticity (omega sub x) is considered first. Two sources are postulated. One is the upstream 'pressure hill', produced by the tab, which appears to be the dominant source. Another is due to vortex filaments shed from the sides of the tab and reoriented downstream by the mean shear of the mixing layer. In the case of a 'delta-tab', a triangular tab with its apex leaning downstream, vorticity from the two sources explain the stronger effect in that configuration. Data on the vorticity evolution for the effect of two delta-tabs are presented, up to twelve jet diameters from the exit, which show that the streamwise vortices persist even at the farthest measurement station. The magnitude of omega sub x-maximum decays continually with distance from the nozzle, its ratio to azimuthal vorticity maximum is found to be about 1/5 everywhere. The relative effect of a delta-tab on jets from an axisymmetric nozzle and a 8:1 rectangular nozzle is also studied. The mixing layer distortion is found to be less pronounced in the rectangular case. The jet mixing, as manifested by the mass flux measured at a downstream station, is increased in the axisymmetric jet but it is decreased in the rectangular jet under consideration by the delta-tab.

Zaman, K. B. M. Q.↗
NASA NTRS2019

Aura to Exit the A-Train? Aura Extended Mission Operations Concept

This presentation provides the future plans for the Earth Observing System (EOS) Aura satellite to the Aura Science Team community. It provides a summary of predictions for future fuel usage, maneuver schemes, and changes to the orbit parameters through the end of mission.

Aura Extended Mission Operations Concept↗
NASA NTRS1987

Unsteady diffuser vane pressure and impeller wake measurements in a centrifugal pump

Unsteady surface pressure measurements on a vaned diffuser of a centrifugal pump, and wake measurements of the flow exiting a centrifugal impeller into a vaneless diffuser are presented. Frequency spectra and ensemble averages are given for the unsteady measurements. Two different impellers were used, the pump impeller of the HPOTP (High Pressure Oxygen Turbopump) of the SSME (Space Shuttle Main Engine) and a two-dimensional impeller. The magnitude of the unsteady total pressure measured in the stationary frame at the impeller exit was found to be of the same order of magnitude as the total pressure rise across the pump. The magnitude of the unsteady diffuser vane pressures was observed to be significantly different on suction and pressure side of the vane, attaining its largest value on the suction side the leading edge while decreasing along the vane.

Arndt, N.↗
NASA NTRS1985

Description and calibration of the Langley Hypersonic CF4 tunnel: A facility for simulating low gamma flow as occurs for a real gas

The Langley Hypersonic CF4 Tunnel is a Mach 6 facility which simulates an important aspect of dissociative real-gas phenomena associated with the reentry of blunt vehicles, i.e., the decrease in the ratio of specific heats (gamma) that occurs within the shock layer of the vehicle. A general description of this facility is presented along with a discussion of the basic components, instrumentation, and operating procedure. Pitot-pressure surveys were made at the nozzle exit and downstream of the exit for reservoir temperatures from 1020 to 1495 R and reservoir pressures from 1000 to 2550 psia. A uniform test core having a diameter of circa 11 in. (0.55 times the nozzle-exit diameter) exists at the maximum value of reservoir pressure and temperature. The corresponding free-stream Mach number is 5.9, the unit Reynolds number is 4 x 10 to the 5th power per foot, the ratio of specific heats immediately behind a normal shock is 1.10, and the normal-shock density ratio is 12.6. When the facility is operated at reservoir temperatures below 1440 R, irregularities occur in the pitot-pressure profile within a small region about the nozzle centerline. These variations in pitot pressure indicate the existence of flow distrubances originating in the upstream region of the nozzle. This necessitates testing models off centerline in the uniform flow between the centerline region and either the nozzle boundary layer or the lip shock originating at the nozzle exit. Samples of data obtained in this facility with various models are presented to illustrate the effect of gamma on flow conditions about the model and the importance of knowing the magnitude of this effect.

Midden, Raymond E.↗
NASA NTRS2016

Numerical Evaluation of an Ejector-Enhanced Resonant Pulse Combustor with a Poppet Inlet Valve and a Converging Exhaust Nozzle

A computational investigation of a pressure-gain combustor system for gas turbine applications is presented. The system consists of a valved pulse combustor and an ejector, housed within a shroud. The study focuses on two enhancements to previous models, related to the valve and ejector components. First, a new poppet inlet valve system is investigated, replacing the previously used reed valve configuration. Secondly, a new computational approach to approximating the effects of choked turbine inlet guide vanes present immediately downstream of the Ejector-Enhanced Resonant Pulse Combustor (EERPC) is investigated. Instead of specifying a back pressure at the EERPC exit boundary (as was done in previous studies) the new model adds a converging-diverging (CD) nozzle at the exit of the EERPC. The throat area of the CD nozzle can be adjusted to obtain the desired back pressure level and total mass flow rate. The results presented indicate that the new poppet valve configuration performs nearly as well as the original reed valve system, and that the addition of the CD nozzle is an effective method to approximate the exit boundary effects of a turbine present downstream of the EERPC. Furthermore, it is shown that the more acoustically reflective boundary imposed by a nozzle as compared to a constant pressure surface does not significantly affect operation or performance.

Computational Fluid Dynamics (CFD↗
NASA NTRS1988

Active cooling design for scramjet engines using optimization methods

A methodology for using optimization in designing metallic cooling jackets for scramjet engines is presented. The optimal design minimizes the required coolant flow rate subject to temperature, mechanical-stress, and thermal-fatigue-life constraints on the cooling-jacket panels, and Mach-number and pressure contraints on the coolant exiting the panel. The analytical basis for the methodology is presented, and results for the optimal design of panels are shown to demonstrate its utility.

Scotti, Stephen J.↗
NASA NTRS1988

Active cooling design for scramjet engines using optimization methods

A methodology for using optimization in designing metallic cooling jackets for scramjet engines is presented. The optimal design minimizes the required coolant flow rate subject to temperature, mechanical-stress, and thermal-fatigue-life constraints on the cooling-jacket panels, and Mach-number and pressure constraints on the coolant exiting the panel. The analytical basis for the methodology is presented, and results for the optimal design of panels are shown to demonstrate its utility.

Scotti, Stephen J.↗
NASA NTRS1993

Plume characteristics of an arcjet thruster

An experimental investigation is conducted of the plume of a 1.5 kW hydrogen-fueled arcjet thruster. The velocity and translational temperature are measured by laser induced fluorescence of excited-state atomic hydrogen at a variety of locations and operating conditions. The peak velocity and temperature at the center of the exit plane of 14.7 km/s and 5000 K lead to a Mach number of 1.9. Evidence is presented which demonstrates the extent to which viscous forces dominate the flow, such as the exit plane Reynolds number of 18. Data taken in the plume suggest a transition from supersonic to subsonic flow via viscous dissipation. In addition, the results suggest that high tank pressures (0.4 to 1.5 Torr) limit the ability to simulate space environments for different portions of the plume.

Liebeskind, John G.↗
NASA NTRS1975

Full scale upper surface blown flap noise

A highly noise suppressed TF 34 engine was used to investigate the noise of several powered lift configurations involving upper surface blown (USB) flaps. The configuration variables were nozzle type (i.e. slot and circular with deflector), flap chord length, and flap angle. The results of velocity surveys at both the nozzle exit and the flap trailing edge are also presented and used for correlation of the noise data. Configurations using a long flap design were 4 db quieter than a short flap typical of current trends in USB flap design. The lower noise for the long flap is attributed primarily to the greater velocity decay of the jet at the flap trailing edge. The full-scale data revealed substantially more quadrupole noise in the region near the deflected jet than observed in previous sub-scale tests.

Heidelberg, L. J.↗
NASA NTRS1977

Aerodynamic characteristics of a small-scale straight and swept-back wing with knee-blown jet flaps

Two sting-mounted, 50.8 cm (20 in.) span, knee-blown, jet-flap models were tested in a large (2.1- by 2.5-m (7- by 10-ft) subsonic wind tunnel. A straight- and swept-wing model were tested with fixed flap deflection with various combinations of full-span leading-edge slats. The swept-wing model was also tested with wing tip extensions. Data were taken at angles-of-attack between 0 deg and 40 deg, at dynamic pressures between 143.6 N/sq m (3 lb/sq ft) and 239.4 N/sq m (5 lb/sq ft), and at Reynolds numbers (based on wing chord) ranging from 100,000 to 132,000. Jet flap momentum blowing coefficients up to 10 were used. Lift, drag, and pitching-moment coefficients, and exit flow profiles for the flap blowing are presented in graphical form without analysis.

Morehouse, G. G.↗
NASA NTRS1979

Investigation of ground effects on large and small scale models of a three fan V/STOL aircraft configuration

Induced lift of a subsonic, three fan, lift/cruise, V/STOL aircraft configuration was investigated using scale modes of a multimission aircraft whose design incorporates a nose mounted lift fan and two lift/cruise units located over the wing. Configuration effects were assessed for lift improvement devices, lift/cruise nozzle rails, nozzle perimeter plates, and alternate nose fan exit hubs. Tests were conducted at four model heights (H/D = 0.95, 1.53, 3.06 and 6.45, where D is the average nozzle exit diameter equal to 0.997 m.) Results are presented and discussed.

Schuster, E. P.↗
NASA NTRS1981

Jet array impingement with crossflow-correlation of streamwise resolved flow and heat transfer distributions

Correlations for heat transfer coefficients for jets of circular offices and impinging on a surface parallel to the jet orifice plate are presented. The air, following impingement, is constrained to exit in a single direction along the channel formed by the jet orifice plate and the heat transfer (impingement) surface. The downstream jets are subjected to a crossflow originating from the upstream jets. Impingement surface heat transfer coefficients resolved to one streamwise jet orifice spacing, averaged across the channel span, are correlated with the associated individual spanwise orifice row jet and crossflow velocities, and with the geometric parameters.

Florschuetz, L. W.↗