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At least 289 records · Page 16
All sourcesNASA NTRSDOE OSTI
NASA NTRS1986

Rotary Speed Sensor for Antilocking Brakes

Sensor based on fluidic principles produces negative pressure approximately proportional to rotational speed. Sensor developed as part of antilocking brake system for motorcycles. Uses inlet pressure rather than outlet pressure as braking-control signal, eliminating pressure pulsations caused by pump vanes and ensuring low-noise signal. Sensor is centrifugal air pump turned by one of motorcycle wheels. Air enters pump through orifice plates, and suction taken off through port in pump inlet plenum.

Berdahl, C. M.↗
NASA NTRS1984

National transonic facility shakedown test results and calibration plans

The results of the shakedown tests and the calibration plan of the National Transonic Facility (NTF) are presented. The facility is designed to operate in both air and nitrogen modes, cover Mach numbers from 0.2 to 1.2, pressures up to 8.8 atm and temperatures between 77 and 339 K. The facility data system is built around four 16-bit minicomputers with a total memory of three megabytes. A portable cryogenic chamber is available. The tunnel systems were operated in a series of tests in Mach number range of 0.2 to 1.17, pressures up to 8.5 atm, and temperatures down to 100 K. The calibration plan includes steady-state and dynamic calibration, as well as wall interference studies. The facility underwent the checkout of the model attitude, plenum isolation, and model access systems, followed by aerodynamic calibration in 1984. Schematic drawings and diagrams are included.

Bruce, W. E., Jr.↗
NASA NTRS1985

On boundary layer transition in high-subsonic and transonic flow under the influence of acoustic disturbances and free-stream turbulence

The existing data base which quantifies the roles played by the Mach number, turbulence level, and acoustic noise in the flow quality of wind tunnels is reviewed. Attention is given to the effects on end (ET) and beginning (BT) transitions in flows over a 10 deg cone in transonic channels. The Re of both ET and BT flows were significantly affected by the Mach number, acoustic noise level and the noise frequency content. A unit increment in the Mach number increased the Re beyond 3,000,000, while noise below the 1 percent level in the rms pressure coefficient did not alter the transition phenomenon. More data are required on the effects of turbulence, which was a significant factor, and in the settling chamber, test section, plenum and diffuser to trace the locations and magnitudes of the acoustic sources.

Murthy, S. V.↗
NASA NTRS1985

Velocity selective flow visualization in a free supersonic nitrogen jet with the resonant Doppler velocimeter

The use of the resonant Doppler velocimeter for flow visualization in a free supersonic nitrogen jet is demonstrated experimentally. In the experiment reported here, room-temperature nitrogen at 12.7 psi expands through a Mach 3.4 converging diverging nozzle into a plenum chamber; small amount of sodium is injected through a heated needle centered 0.5-in. upstream of the nozzle throat. The beam of a single frequency dye laser is expanded into a sheet of light with a cylindrical beam expander and directed counter to the jet to spatially resolve any slice across the core of the flow. By tuning the laser to different frequencies, particular velocity and density fields can be highlighted.

Zimmermann, M.↗
NASA NTRS1985

Particle generation experience in Langley's 16-foot transonic tunnel

The Langley 16-Foot Transonic Tunnel is a single-return, continuous-flow, atmospheric tunnel which uses air exchange for cooling. The tunnel speed is continuously variable from Mach 0 to 1.30. The test section is a regular octagon in cross-section with slots at the corners of the octagon. The laser velocimeter optics, laser, photomultipliers, etc. are mounted in the test section plenum chamber which surrounds the test section. The test volume is approximately a 1-meter cube about the tunnel center line centered on tunnel station 40.84 meters. The seeding system particle generators are mounted on the upstream side of the fourth set of turning vanes approximately 49 meters upstream of the test volume. At this point the tunnel has a diameter of 17.68 meters which meant that installation of the generators was a difficult procedure. It might be noted that particle generation had to be continuous as when the generators were turned off the data rate rapidly deteriorated to zero.

Reubush, D. E.↗
NASA NTRS1986

Turbulent dispersion of the icing cloud from spray nozzles used in icing tunnels

To correctly simulate flight in natural icing conditions, the turbulence in an icing simulator must be as low as possible. But some turbulence is required to mix the droplets from the spray nozzles and achieve an icing cloud of uniform liquid water content. The goal for any spray system is to obtain the widest possible spray cloud with the lowest possible turbulence in the test section of a icing tunnel. This investigation reports the measurement of turbulence and the three-dimensional spread of the cloud from a single spray nozzle. The task was to determine how the air turbulence and cloud width are affected by spray bars of quite different drag coefficients, by changes in the turbulence upstream of the spray, the droplet size, and the atomizing air. An ice accretion grid, located 6.3 m downstream of the single spray nozzle, was used to measure cloud spread. Both the spray bar and the grid were located in the constant velocity test section. Three spray bar shapes were tested: the short blunt spray bar used in the NASA Lewis Icing Research Tunnel, a thin 14.6 cm chord airfoil, and a 53 cm chord NACA 0012 airfoil. At the low airspeed (56 km/hr) the ice accretion pattern was axisymmetric and was not affected by the shape of the spray bar. At the high airspeed (169 km/hr) the spread was 30 percent smaller than at the low airspeed. For the widest cloud the spray bars should be located as far upstream in the low velocity plenum of the icing tunnel. Good comparison is obtained between the cloud spread data and predicitons from a two-dimensional cloud mixing computer code using the two equation turbulence (k epsilon g) model.

Marek, C. J.↗
NASA NTRS1985

Model deformation measurements at a cryogenic wind tunnel using photogrammetry

A photogrammetric closed circuit television system to measure model deformation at the National Transonic Facility (NTF) is described. The photogrammetric approach was chosen because of its inherent rapid data recording of the entire object field. Video cameras are used to acquire data instead of film cameras due to the inaccessibility of cameras which must be housed within the cryogenic, high pressure plenum of this facility. Data reduction procedures and the results of tunnel tests at the NTF are presented.

Burner, A. W.↗
NASA NTRS1987

Noise generated by flow through large butterfly valves

A large butterfly valve (1.37 m diam) was acoustically tested to measure the noise generated and propagating in both the upstream and downstream directions. The experimental investigation used wall mounted pressure transducers to measure the fluctuating component of the pipe static pressure upstream and downstream of the valve. Microphones upstream of the pipe inlet and located in a plenum were used to measure the noise radiated from the valve in the upstream direction. Comparison of the wall pressure downstream of the valve to a prediction were made. Reasonable agreement was obtained with the valve operating at a choked condition. The noise upstream of the valve is 30 dB less than that measured downstream.

Huff, Ronald G.↗
NASA NTRS1986

Computation of nozzle flow fields using the PARC2D Navier-Stokes code

Supersonic nozzles which operate at low Reynolds numbers and have large expansion ratios have very thick boundary layers at their exit. This leads to a very strong viscous/inviscid interaction upon the flow within the nozzle and the traditional nozzle design techniques which correct the inviscid core with a boundary layer displacement do not accurately predict the nozzle exit conditions. A full Navier-Stokes code (PARC2D) was used to compute the nozzle flow field. Grids were generated using the interactive grid generator code TBGG. All computations were made on the NASA MSFC CRAY X-MP computer. Comparison was made between the computations and in-house wall pressure measurements for CO2 flow through a conical nozzle having an area ratio of 40. Satisfactory agreement existed between the computations and measurements for a stagnation pressure of 29.4 psia and stagnation temperature of 1060 R. However, agreement did not exist at a stagnation pressure of 7.4 psia. Several reasons for the lack of agreement are possible. The computational code assumed a constant gas gamma whereas gamma for CO2 varied from 1.22 in the plenum chamber to 1.38 at the nozzle exit. Finally, it is possible that condensation occurred during the expansion at the lower stagnation pressure.

Collins, Frank G.↗
NASA NTRS1986

Experimental evaluation of wall Mach number distributions of the octagonal test section proposed for NASA Lewis Research Center's altitude wind tunnel

Wall Mach number distributions were determined over a range of test-section free-stream Mach numbers from 0.2 to 0.92. The test section was slotted and had a nominal porosity of 11 percent. Reentry flaps located at the test-section exit were varied from 0 (fully closed) to 9 (fully open) degrees. Flow was bled through the test-section slots by means of a plenum evacuation system (PES) and varied from 0 to 3 percent of tunnel flow. Variations in reentry flap angle or PES flow rate had little or no effect on the Mach number distributions in the first 70 percent of the test section. However, in the aft region of the test section, flap angle and PES flow rate had a major impact on the Mach number distributions. Optimum PES flow rates were nominally 2 to 2.5 percent wtih the flaps fully closed and less than 1 percent when the flaps were fully open. The standard deviation of the test-section wall Mach numbers at the optimum PES flow rates was 0.003 or less.

Harrington, Douglas E.↗
NASA NTRS1986

The evolution of adaptive-wall wind tunnels

Since its inception, adaptive wall wind tunnel design has developed three major configurational possibilities: streamlined walls, variable porosity walls, and segmented plenum arrangements. All of these methods have demonstrated their feasibility for transonic flow research; development programs are noted to be underway for three-dimensional adaptive wall tunnels. With the availability of advanced computational procedures, routine angle-of-incidence or Mach sweeps may be treated numerically, while adaptive wall wind tunnels are employed for complementary point-design and for testing at flow conditions that are beyond the capabilities of current numerical modeling.

Davis, S. S.↗
NASA NTRS1986

Construction of a 2- by 2-foot transonic adaptive-wall test section at the NASA Ames Research Center

The development of a new production-size, two-dimensional, adaptive-wall test section with ventilated walls at the NASA Ames Research Center is described. The new facility incorporates rapid closed-loop operation, computer/sensor integration, and on-line interference assessment and wall corrections. Air flow through the test section is controlled by a series of plenum compartments and three-way slide vales. A fast-scan laser velocimeter was built to measure velocity boundary conditions for the interference assessment scheme. A 15.2-cm- (6.0-in.-) chord NACA 0012 airfoil model will be used in the first experiments during calibration of the facility.

Morgan, Daniel G.↗
NASA NTRS1987

Navier-Stokes simulation of real gas flows in nozzles

Air flow in a hypersonic nozzle causes real gas effects due to reaction among the species constituting air. Such reactions may be in chemical equilibrium or in chemical nonequilibrium. Here using the CSCM upwind scheme for the compressible Navier-Stokes equations, the real gas flowfield in an arcjet nozzle is computed for both the equilibrium case and the nonequilibrium case. A hypersonic nozzle flow arising from a pebble bed heated plenum is also computed for the equilibrium situation. Between the equilibrium cases, the chemistry is treated by two different schemes and comments are made as to computational complexity. For the nonequilibrium case, a full set of seventeen reactions and full implicit coupling of five species with gasdynamics is employed to compute the flowfield. For all cases considered here the gas is assumed to be a calorically imperfect mixture of ideal gases in thermal equilibrium.

Nagaraj, N.↗
NASA NTRS1987

An 8-cm ion thruster characterization

The performance of the Ion Auxiliary Propulsion System (IAPS) thruster was increased to thrust T = 32 mN, specific impulse I sub sp = 4062 s, and thrust-to-power ratio T/P = 33 mN/kW. This performance was obtained by increasing the discharge power, accelerating voltage, propellant flow rate, and chamber magnetic field. Adding a plenum and main vaporizer for propellant distribution was the only major change required in the thruster. The modified thruster characterization is presented. A cathode magnet assembly did not improve performance. A simplified power processing unit was designed and evaluated. This unit decreased the parts count of the IAPS power processing unit by a factor of ten.

Wessel, F. J.↗
NASA NTRS1987

Low thrust viscous nozzle flow fields prediction

An existing Navier-Stokes code (PARC2D) was used to compute the nozzle flow field. Grids were generated by the interactive grid generator codes TBGG and GENIE. All computations were made on the NASA/MSFC CRAY X-MP computer. Comparisons were made between the computations and MSFC in-house wall pressure measurements for CO2 flow through a conical nozzle having an area ratio of 40. Satisfactory agreements exist between the computations and measurements for different stagnation pressures of 29.4, 14.7, and 7.4 psia, at stagnation temperature of 1060 R. However, agreements did not match precisely near the nozzle exit. Several reasons for the lack of agreement are possible. The computational code assumes a constant gas gamma, whereas the gamma i.e. the specific heat ratio for CO2 varied from 1.22 in the plenum chamber to 1.38 at the nozzle exit. The computations also assumes adiabatic and no-slip walls. Both assumptions may not be correct. Finally, it is possible that condensation occurs during the nozzle expansion at the low stagnation pressure. The next phase of the work will incorporate variable gamma and slip wall boundary conditions in the computational code and develop a more accurate computer code.

Liaw, Goang-Shin↗
NASA NTRS1986

Preliminary calibration and test results from the National Transonic Facility

The National Transonic Facility (NTF) was operated to design condition of 120 million Reynolds number at a Mach number of 1.0. All systems were checked out except plenum isolation valves; modifications are being made to heaters on the actuators. Initial steady-state calibration indicates excellent steady flow characteristics. The first test of the Pathfinder 1 model indicated significant Reynolds number effects. Some effect of temperature on instrumentation were obtained. The cause of these effects is being evaluated.

Mckinney, Linwood W.↗
NASA NTRS1987

The impact of circulation control on rotary aircraft controls systems

Application of circulation to rotary wing systems is a new development. Efforts to determine the near and far field flow patterns and to analytically predict those flow patterns have been underway for some years. Rotary wing applications present a new set of challenges in circulation control technology. Rotary wing sections must accommodate substantial Mach number, free stream dynamic pressure and section angle of attack variation at each flight condition within the design envelope. They must also be capable of short term circulation blowing modulation to produce control moments and vibration alleviation in addition to a lift augmentation function. Control system design must provide this primary control moment, vibration alleviation and lift augmentation function. To accomplish this, one must simultaneously control the compressed air source and its distribution. The control law algorithm must therefore address the compressor as the air source, the plenum as the air pressure storage and the pneumatic flow gates or valves that distribute and meter the stored pressure to the rotating blades. Also, mechanical collective blade pitch, rotor shaft angle of attack and engine power control must be maintained.

Kingloff, R. F.↗
NASA NTRS1988

Initial condition effect on pressure waves in an axisymmetric jet

A pair of microphones (separated axially by 5.08 cm and laterally by 1.3 cm) are placed on either side of the jet centerline to investigate coherent pressure fluctuations in an axisymmetric jet at Strouhal numbers less than unity. Auto-spectra, transfer-function, and coherence measurements are made for a tripped and untripped boundary layer initial condition. It was found that coherent acoustic pressure waves originating in the upstream plenum chamber propagate a greater distance downstream for the tripped initial condition than for the untripped initial condition. In addition, for the untripped initial condition the development of the coherent hydrodynamic pressure waves shifts downstream.

Miles, Jeffrey H.↗