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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 109 records · Page 6
All sourcesNASA NTRSDOE OSTI
NASA NTRS2001

Subsonic Glideback Rocket Demonstrator Flight Testing

For the past two years, Cal Poly's rocket program has been aggressively exploring the concept of remotely controlled, fixed wing, flyable rocket boosters. This program, embodied by a group of student engineers known as Cal Poly Space Systems, has successfully demonstrated the idea of a rocket design that incorporates a vertical launch pattern followed by a horizontal return flight and landing. Though the design is meant for supersonic flight, CPSS demonstrators are deployed at a subsonic speed. Many steps have been taken by the club that allowed the evolution of the StarBooster prototype to reach its current size: a ten-foot tall, one-foot diameter, composite material rocket. Progress is currently being made that involves multiple boosters along with a second stage, third rocket.

DeTurris, Dianne J.↗
NASA NTRS1983

An experimental documentation of trailing-edge flows at high Reynolds number

Experiments documenting attached trailing-edge and near-wake flows at high Reynolds numbers are described. A long, airfoil-like model was tested at subsonic and low transonic Mach numbers, and both symmetrical and asymmetrical flows with pressure gradients upstream of the trailing edge were investigated. Model surface pressures and detailed mean and turbulence flow qualities were measured in the vicinity of the trailing edge and in the near-wake. The data obtained are of sufficient quality and detail to be useful as test cases in assessing turbulence models and calculation methods.

Viswanath, P. R.↗
NASA NTRS1990

Impact of nose-probe chines on the vortex flows about the F-16C

A 1/15-scale model of the F-16C aircraft was tested at subsonic and transonic speeds to determine the effects of small nose-probe chines on the high angle-of-attack aerodynamic and stability characteristics. Off-body flow visualization using a laser vapor screen technique and six-component forces and moments were obtained at freestream Mach numbers of 0.30, 0.60, and 0.80. The nose-probe chines generated strong vortices interacting with the vortices generated over the smooth forebody or over the strakes. These vortices in turn led to global flow unsteadiness and model dynamics at the high angles of attack. The test results underscore the sensitivity of the high angle-of-attack flow field to small geometry changes in the nose region.

Hall, R. M.↗
NASA NTRS1959

Some Static Oscillatory and Free Body Tests of Blunt Bodies at Low Subsonic Speeds

Some blunt-body shapes considered suitable for entry into the earth's atmosphere were tested by both static and oscillatory methods in the Langley stability tunnel. In addition, free-fall tests of some similar models were made in the Langley 20-foot free-spinning tunnel. The results of the tests show that increasing the flare of the body shape increased the dynamic stability and that for flat-faced shapes increasing the corner radius increased the stability. The test data from the Langley stability tunnel were used to compute the damping factor for the models tested in the langley 20-foot free-spinning tunnel. For these cases in which the damping factor was low, -1/2 or less, the stability was critical and sensitive to disturbance. When the damping factor was about -2, damping was generally obtained.

Lichtenstein, Jacob H.↗
NASA NTRS1982

Energy efficient engine: Low-pressure turbine subsonic cascade component development and integration program

A subsonic cascade test program was conducted to provide technical data for optimizing the blade and vane airfoil designs for the Energy Efficient Engine Low-Pressure Turbine component. The program consisted of three parts. The first involved an evaluation of the low-chamber inlet guide vane. The second, was an evaluation of two candidate aerodynamic loading philosophies for the fourth blade root section. The third part consisted of an evaluation of three candidate airfoil geometries for the fourth blade mean section. The performance of each candidate airfoil was evaluated in a linear cascade configuration. The overall results of this study indicate that the aft-loaded airfoil designs resulted in lower losses which substantiated Pratt & Whitney Aircraft's design philosophy for the Energy Efficient Engine low-pressure turbine component.

Sharma, O. P.↗
NASA NTRS1987

Analysis of 7- X 10-foot high speed wind tunnel shaft loads in support of fan blade failure investigation

This is a report of the investigation of the High-Speed 7- X 10-Foot Wind Tunnel at NASA Langley Research Center, which experienced a catastrophic failure of all 18 Sitka spruce fan blades during operation at 0.8 Mach number on 2 July 1985. The High-Speed Tunnel, a closed-circuit/single-return atmospheric wind tunnel, had been operated since 1945 to support a wide range of subsonic aerodynamic tests and studies. The failed blade set had been in use since 1975. In addition to blade loss, the most significant damage was a bent main drive shaft for a total estimated damage loss of 1.7 million dollars. An analysis of the natural frequency characteristics as well as loads, reactions, stresses, and deflections of the fan drive system resulting from steady-state and dynamic loads due to unbalance was performed. Transient load cases were simulated by step input and ramp input loading functions intended to simulate the loss of one to nine blades (maximum unbalance forces).

Faison, Richard W.↗
NASA NTRS1989

Experimental transition and boundary-layer stability analysis for a slotted swept laminar flow control airfoil

A swept, supercritical laminar flow control (LFC) airfoil designated NASA SCLFC(1)-0513F was tested at subsonic and transonic speeds in the NASA Langley eight-foot Transonic Pressure Tunnel. This paper examines Tollmien-Schlichting and crossflow disturbance amplification for this airfoil using the linear stability method. The design methodology using linear stability analysis is evaluated and the results of the incompressible and compressible methods are compared. Experimental data on the swept, supercritical LFC airfoil and reference wind tunnel and flight results are used to correlate and evaluate the N-factor method for transition prediction over a speed range M(infinity) from zero to one.

Harvey, William D.↗
NASA NTRS1992

Experimental investigation of the effects of aft blowing with various nozzle exit geometries on a 3.0 caliber tangent ogive at high angles of attack: Forebody pressure distributions

An experimental study of the effects of aft blowing on the asymmetric vortex flow of a slender, axisymmetric body at high angles of attack was conducted. A 3.0 caliber tangent ogive body fitted with a cylindrical afterbody was tested in a wind tunnel under subsonic, laminar flow test conditions. Asymmetric blowing from both a single nozzle and a double nozzle configuration, positioned near the body apex, was studied. Aft blowing was observed to alter the vortex asymmetry by moving the blowing-side vortex closer to the body surface while moving the non-blowing-side vortex further away from the body. The effect of increasing the blowing coefficient was to move the blowing-side vortex closer to the body surface at a more upstream location. The data also showed that blowing was more effective in altering the initial vortex asymmetry at the higher angles of attack than at the lower. The effects of changing the nozzle exit geometry were studied and it was observed that blowing from a nozzle with a low, broad exit geometry was more effective in reducing the vortex asymmetry than blowing from a high, narrow exit geometry.

Chokani, Ndaona↗
NASA NTRS1992

An investigation of the effects of aft blowing on a 3.0 caliber tangent ogive body at high angles of attack

An experimental investigation of the effects of aft blowing on the asymmetric vortex flow of a slender, axisymmetric body at high angles of attack was conducted. A 3.0 caliber tangent ogive body fitted with a cylindrical afterbody was tested in a wind tunnel under subsonic, laminar flow test conditions. Asymmetric blowing from both a single nozzle and a double nozzle configuration, positioned near the body apex, was investigated. Aft blowing was observed to alter the vortex asymmetry by moving the blowing-side vortex closer to the body surface while moving the non-blowing-side vortex further away from the body. The effect of increasing the blowing coefficient was to move the blowing-side vortex closer to the body surface at a more upstream location. The data also showed that blowing was more effective in altering the initial vortex asymmetry at the higher angles of attack than at the lower. The effects of changing the nozzle exit geometry were investigated and it was observed that blowing from a nozzle with a low, broad exit geometry was more effective in reducing the vortex asymmetry than blowing from a high, narrow exit geometry.

Gittner, Nathan M.↗
NASA NTRS1941

Temperature recording in high-speed gases

The development of suitable thermometer shapes giving the amount of temperature rise if possible without calibration and affording ready repetition is predicated upon a fundamental elucidation of this heating on a number of elementary body forms. This report provides results of tests at subsonic and supersonic velocities.

Eckert, E↗
NASA NTRS2001

Large Engine Technology (LET) Task XXXVII Low-Bypass Ratio Mixed Turbofan Engine Subsonic Jet Noise Reduction Program Test Report

NASA Glenn Research Center supported a three year effort to develop the technology for reducing jet noise from low-bypass ratio engines. This effort concentrated on both analytical and experimental approaches using various mixer designs. CFD and MGB predictions are compared with LDV and noise data, respectively. While former predictions matched well with data, experiment shows a need for improving the latter predictions. Data also show that mixing noise can be sensitive to engine hardware upstream of the mixing exit plane.

Hauser, Joseph R.↗
NASA NTRS2003

StarBooster Demonstrator Cluster Configuration Analysis/Verification Program

In order to study the flight dynamics of the cluster configuration of two first stage boosters and upper-stage, flight-testing of subsonic sub-scale models has been undertaken using two glideback boosters launched on a center upper-stage. Three high power rockets clustered together were built and flown to demonstrate vertical launch, separation and horizontal recovery of the boosters. Although the boosters fly to conventional aircraft landing, the centerstage comes down separately under its own parachute. The goal of the project has been to collect data during separation and flight for comparison with a six degree of freedom simulation. The configuration for the delta wing canard boosters comes from a design by Starcraft Boosters, Inc. The subscale rockets were constructed of foam covered in carbon or fiberglass and were launched with commercially available solid rocket motors. The first set of boosters built were 3-ft tall with a 4-ft tall centerstage, and two additional sets of boosters were made that were each over 5-ft tall with a 7.5 ft centerstage. The rocket cluster is launched vertically, then after motor bum out the boosters are separated and flown to a horizontal landing under radio-control. An on-board data acquisition system recorded data during both the launch and glide phases of flight.

DeTurris, Dianne J.↗
NASA NTRS2005

Stagnation Point Heat Transfer with Gas Injection Cooling

The present paper deals with an experimental study of the stagnation-point heat transfer to a cooled copper surface with gas injection under subsonic conditions. Test were made with a probe that combined a steady-state water-cooled calorimeter that allows the capability to study convective blockage and to perform heat transfer measurements in presence of gas injection in the stagnation region. The copper probe was pierced by 52 holes, representing 2.4% of the total probe surface. The 1.2 MW high enthalpy plasma wind tunnel was operated at anode powers between 130 and 230 kW and a static pressures from 35 hPa up to 200 hPa. Air, carbon dioxide and argon were injected in the mass flow range 0-0.4 g/s in the boundary layer developed around the 50 mm diameter probe. The measured stagnation-point heat transfer rates are reported and discussed.

Vancrayenest, B.↗
NASA NTRS1950

An Investigation of the Longitudinal Characteristics of the MX-656 Configuration Using Rocket-Propelled Models Preliminary Results at Mach Numbers from 0.65 to 1.25

A rocket-propelled model of the Mx-656 configuration has been flown through the Mach number range from 0.65 to 1.25. An analysis of the response of the model to rapid deflections of the horizontal tail gave information on the lift, drag, longitudinal stability and control, and longitudinal-trim change. The lift-coefficient range covered by the test was from -0.2 to 0,3 throughout most of the Mach number range, The model was statically and dynamically stable throughout the lift-coefficient and Mach number range of the test. At subsonic speeds the aerodynamic center moved f o m r d with increasing lift coefficient. The most forward position of the aerodynamic center was about 12,5 percent of the mean aerodynamic chord at a small positive lift coefficient and at a Mach number of about 0.84. A t supersonic speeds the aerodynamic center was well aft, varying from 33 to 39 percent of the mean aerodynamic chord at Mach numbers of 1.0 and 1.25, respectively. Transonic-trim change, as measured by the change in trim lift coefficient with Mach number at a constant t a i l setting, was of small magnitude (about 0.1 lift coefficient for zero tail setting). The zero-lift/drag coefficient increased about 0.042 in the region between a Mach number of 0.9 and 1.1

Mitchell, Jesse L.↗
NASA NTRS2019

The Effect of Incoming Boundary Layer Characteristics on the Performance of a Distributed Propulsion System

An experimental study is conducted in an effort to advance the understanding of flow physics associated with a boundary layer ingesting, distributed propulsion system. The influence of incoming boundary layer thickness and flow distortions are studied on the flow downstream and the overall performance of the system. The propulsion model, fabricated using additive manufacturing and integrated with electrical fans, is mounted on a flat plate and tested at subsonic speeds. Detailed characterization of the incoming boundary layer and subsequent assessment of the downstream flow field is performed using hotwire anemometry. Modification of the incoming boundary layer is achieved by placing tripping devices, such as rods and vortex generating ramps, near the leading edge of the flat plate. The overall performance of the system for different incoming flow conditions is analyzed by comparing magnitudes of exhaust velocities as well as estimated propulsive power to the corresponding baseline values. For a constant input power to the fans, smaller upstream flow distortions and moderately thickened boundary layers result in marginal changes in the flow field downstream. On the other hand, notable reductions in downstream flow velocities and propulsive power are observed in the case of a significantly thicker and/or distorted incoming boundary layer. It is hoped that this study will serve as a database for this technologically relevant flow field that has not been explored adequately before.

Propulsion↗