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At least 451 records · Page 25

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NASA NTRS1994

Integrated Tail Buffet Loads on the F/A-18

The unsteady pressures acting on the vertical tails of a full-scale F/A-18 fighter aircraft were studied to gain a better understanding of tail-buffet loads that frequently occur on fighter aircraft operating at high angles-of-attack. Data for the study were acquired during two test entries in the 80- by 120-Foot Wind Tunnel wherein the aircraft was tested at wind speeds up to 100 knots and at angles-of-attack from 20deg to 40deg. For the purposes of this paper, the primary difference between the two tests is that, during the first wind-tunnel entry, the pressure transducers were more sparsely spaced and covered less of the fin than during the second entry. In addition to a brief description of the spectral analysis methods used for the unsteady aerodynamic pressures and loads, an analysis of the effects of sensor density on estimating integrated loads is presented. It was found that the integrated loads determined from sparse sensor arrays are significantly higher than actual loads. However, a modest increase in the number of sensors can greatly reduce the error and a method for correcting load estimates from sparse sensor arrays is also suggested. The results for the time-averaged, power-spectral analysis are then presented for the tail-fin bending moments. Power spectra are presented for the aircraft at zero sideslip over an angle-of-attack range from 20deg to 40deg and for the aircraft at an angle-of-attack of 30deg over a sideslip range from -16deg to 16deg. Since the aircraft was equipped with a removable leading-edge extension (LEX) fence to reduce tail-buffet loads, the tail-fin bending moment loads are also presented for that configuration. The LEX fence is shown to significantly reduce bending moment loads over a broad range frequencies, for all the aircraft attitudes presented.

Meyn, Larry A.↗

NASA NTRS1978

Dynamic stall experiments on the NACA 0012 airfoil

The flow over a NACA 0012 airfoil undergoing large oscillations in pitch was experimentally studied at a Reynolds number of and over a range of frequencies and amplitudes. Hot-wire probes and surface-pressure transducers were used to clarify the role of the laminar separation bubble, to delineate the growth and shedding of the stall vortex, and to quantify the resultant aerodynamic loads. In addition to the pressure distributions and normal force and pitching moment data that have often been obtained in previous investigations, estimates of the unsteady drag force during dynamic stall have been derived from the surface pressure measurements. Special characteristics of the pressure response, which are symptomatic of the occurrence and relative severity of moment stall, have also been examined.

Mcalister, K. W.↗

NASA NTRS2012

Deployment of a Pressure Sensitive Paint System for Measuring Global Surface Pressures on Rotorcraft Blades in Simulated Forward Flight

This paper will present details of a Pressure Sensitive Paint (PSP) system for measuring global surface pressures on the tips of rotorcraft blades in simulated forward flight at the 14- x 22-Foot Subsonic Tunnel at the NASA Langley Research Center. The system was designed to use a pulsed laser as an excitation source and PSP data was collected using the lifetime-based approach. With the higher intensity of the laser, this allowed PSP images to be acquired during a single laser pulse, resulting in the collection of crisp images that can be used to determine blade pressure at a specific instant in time. This is extremely important in rotorcraft applications as the blades experience dramatically different flow fields depending on their position in the rotor disk. Testing of the system was performed using the U.S. Army General Rotor Model System equipped with four identical blades. Two of the blades were instrumented with pressure transducers to allow for comparison of the results obtained from the PSP. Preliminary results show that the PSP agrees both qualitatively and quantitatively with both the expected results as well as with the pressure taps. Several areas of improvement have been indentified and are currently being developed.

Watkins, A. Neal↗

NASA NTRS1990

An improved method for simultaneous determination of frictional pressure drop and vapor volume fraction in vertical flow boiling

The two-phase frictional pressure drop and vapor volume fraction in the vertical boiling and adiabatic flow of the refrigerant, R11, have been simultaneously measured by a liquid balancing column and differential magnetic reluctance pressure transducers. An account is given of the experimental apparatus and procedure, data acquisition and analysis, and error estimation employed. All values of two-phase multipliers evaluated on the basis of the measured frictional pressure drop data in vertical upflow fall in the range bounded by the predictions of the Chisholm correlation and the homogeneous model.

Klausner, J. F.↗

NASA NTRS1991

Analysis of Fluctuating Static Pressure Measurements in a Large High Reynolds Number Transonic Cryogenic Wind Tunnel

Dynamic measurements of fluctuating static pressure levels were made using flush mounted high frequency response pressure transducers at eleven locations in the circuit of the National Transonic Facility (NTF) over the complete operating range of this wind tunnel. Measurements were made at test section Mach numbers from 0.2 to 1.2, at pressure from 1 to 8.6 atmospheres and at temperatures from ambient to -250 F, resulting in dynamic flow disturbance measurements at the highest Reynolds numbers available in a transonic ground test facility. Tests were also made independently at variable Mach number, variable Reynolds number, and variable drivepower, each time keeping the other two variables constant thus allowing for the first time, a distinct separation of these three important variables. A description of the NTF emphasizing its flow quality features, details on the calibration of the instrumentation, results of measurements with the test section slots covered, downstream choke, effects of liquid nitrogen injection and gaseous nitrogen venting, comparisons between air and nitrogen, isolation of the effects of Mach number, Reynolds number, and fan drive power, and identification of the sources of significant flow disturbances is included. The results indicate that primary sources of flow disturbance in the NTF may be edge-tones generated by test section sidewall re-entry flaps and the venting of nitrogen gas from the return leg of the tunnel circuit between turns 3 and 4 in the cryogenic mode of operation. The tests to isolate the effects of Mach number, Reynolds number, and drive power indicate that Mach number effects predominate. A comparison with other transonic wind tunnels shows that the NTF has low levels of test section fluctuating static pressure especially in the high subsonic Mach number range from 0.7 to 0.9.

Igoe, William B.↗

NASA NTRS1996

Analysis of Fluctuating Static Pressure Measurements in the National Transonic Facility

Dynamic measurements of fluctuating static pressure levels were taken with flush-mounted, high-frequency response pressure transducers at 11 locations in the circuit of the National Transonic Facility (NTF) across the complete operating range of this wind tunnel. Measurements were taken at test-section Mach numbers from 0.1 to 1.2, at pressures from 1 to 8.6 atm, and at temperatures from ambient to -250 F, which resulted in dynamic flow disturbance measurements at the highest Reynolds numbers available in a transonic ground test facility. Tests were also made by independent variation of the Mach number, the Reynolds number, or the fan drive power while the other two parameters were held constant, which for the first time resulted in a distinct separation of the effects of these three important parameters.

Igoe, William B.↗

NASA NTRSDate not supplied

Analysis of Transonic Unsteady Aerodynamic Environments using Unsteady Pressure Sensitive Paint for the Space Launch System Block 1 Cargo Launch Vehicle

Predicting launch vehicle unsteady aerodynamic loads due to buffet remains a significant challenge. Current practices for modeling buffet environments involve the development of buffet forcing functions using unsteady pressure measurements acquired during wind-tunnel tests. These practices often result in significant uncertainty in buffet environments for coupled loads analyses due to the complex spatio-temporal nature of the unsteady pressure field and challenge of its estimation using discrete sensors. Unsteady pressure sensitive paint, on the other hand, can provide unsteady pressure data at a comparatively high-spatial-density and may overcome the challenge of unsteady pressure field estimation with discrete sensors and lead to improvements in the development of buffet forcing functions. In this paper, comparisons of the fluctuating pressure field are made for the Space Launch System Block 1 cargo launch vehicle measured using unsteady pressure sensitive paint and pressure transducers.

buffet↗

NASA NTRSDate not supplied

Analysis of Transonic Unsteady Aerodynamic Environments using Unsteady Pressure Sensitive Paint for the Space Launch System Block 1 Cargo Launch Vehicle

Predicting launch vehicle unsteady aerodynamic loads due to buffet remains a significant challenge. Current practices for modeling buffet environments involve the development of buffet forcing functions using discrete unsteady pressure measurements acquired during wind-tunnel tests. These practices often result in significant uncertainty in buffet environments for coupled loads analyses due to the complex spatio-temporal nature of the unsteady pressure field and the challenge of its estimation using discrete sensors. Unsteady pressure sensitive paint, on the other hand, can provide unsteady pressure data at a comparatively high spatial density and may overcome the challenge of unsteady pressure field estimation with discrete sensors and lead to improvements in the development of buffet forcing functions. In this paper, comparisons of the fluctuating pressure field are made for the Space Launch System Block 1 cargo launch vehicle measured using unsteady pressure sensitive paint and pressure transducers.

buffet↗

NASA NTRSDate not supplied

Analysis of Transonic Unsteady Aerodynamic Environments using Unsteady Pressure Sensitive Paint for the Space Launch System Block 1 Cargo Launch Vehicle

Predicting launch vehicle unsteady aerodynamic loads due to buffet remains a significant challenge. Current practices for modeling buffet environments involve the development of buffet forcing functions using discrete unsteady pressure measurements acquired during wind-tunnel tests. These practices often result in significant uncertainty in buffet environments for coupled loads analyses due to the complex spatio-temporal nature of the unsteady pressure field and the challenge of its estimation using discrete sensors. Unsteady pressure sensitive paint, on the other hand, can provide unsteady pressure data at a comparatively high spatial density and may overcome the challenge of unsteady pressure field estimation with discrete sensors and lead to improvements in the development of buffet forcing functions. In this paper, comparisons of the fluctuating pressure field are made for the Space Launch System Block 1 cargo launch vehicle measured using unsteady pressure sensitive paint and pressure transducers.

buffet↗

NASA NTRS1983

Feasibility study of three-dimensional holographic interferometry for aerodynamics

Laser holographic interferometry was tried as a nonintrusive diagnostic tool for studying unsteady two dimensional flows. A NACA 0012 airfoil was tested, while undergoing dynamic stall, over a range of Mach numbers of 0.3 to 0.5, Reynolds number of 0.5 x 10 to the 6th power to 2 x 10 to the 6th power, and at reduced frequencies of 0.015 to 0.15. It was found that both quantitative and qualitative data could be obtained by the technique. Surface pressures on the airfoil can be measured to within 1% of those measured with orifices and pressure transducers when the flow is attached. Velocity profiles were measured near the wake region, and they compared very well with laser velocimeter data for attached flows. For separated flows with large scale vortices, densities can be measured, but pressures and velocities cannot be deduced with the assumption of constant pressure gradient in the normal direction. The sensitivity of the interferograms was good at a Mach number of 0.4 and a Reynolds number of 4 x 10 to the 6th power/ft; the sensitivity worsened at smaller Mach numbers and Reynolds numbers, and improved at larger ones.

Craig, J. E.↗

NASA NTRS1993

Modelling, calibration, and error analysis of seven-hole pressure probes

This report describes the calibration of a nonnulling, conical, seven-hole pressure probe over a large range of flow onset angles. The calibration procedure is based on the use of differential pressures to determine the three components of velocity. The method allows determination of the flow angle and velocity magnitude to within an average error of 1.0 deg and 1.0 percent, respectively. Greater accuracy can be achieved by using high-quality pressure transducers. Also included is an examination of the factors which limit the use of the probe, a description of the measurement chain, an error analysis, and a typical experimental result. In addition, a new general analytical model of pressure probe behavior is described, and the validity of the model is demonstrated by comparing it with experimentally measured calibration data for a three-hole yaw meter and a seven-hole probe.

Zillac, G. G.↗

NASA NTRS1972

Effects of wind-tunnel disturbances on hypersonic boundary-layer transition. I, II.

An experimental investigation of the dependence on facility disturbance of transition on sharp cones has been made. In Part I, the sound radiated from the turbulent boundary layer on the walls of several facilities was measured with pressure transducers mounted flush with the cone surface. In Part II, a constant current hot-wire anemometer measured free-stream and cone shock layer disturbances in two hypersonic helium tunnels. Comparison of hot-wire results with surface pressure measurements in one of the helium tunnels indicate the latter data provide an accurate indication of the facility disturbance levels. Both studies show that transition Reynolds numbers correlate in terms of facility rms sound disturbance levels provided the laminar boundary layers on the models are similar.

Stainback, P. S.↗

NASA NTRS1994

High-Temperature Microphone With Fiber-Optic Output

Acoustic-pressure transducer (microphone) with fiber-optic output designed to withstand hot, loud, structurally vibrating environment like that of jet engine. Features flat frequency response out to frequencies well beyond several-kilohertz range needed to test for acoustic-pressure loads on engine structures.

Hellbaum, Richard F.↗

NASA NTRS1979

Aircraft wake flow effect and horizontal tail buffet

As part of a program to investigate the fluctuating pressure distribution and response behavior of a fighter aircraft in transonic maneuver, an F-5A scale model has previously been tested in an 11-ft transonic wind tunnel. The model, with a number of static and dynamic pressure transducers imbedded in the lifting surfaces was tested at various angles of attack up to 16 deg. In this paper, test results of particular interest to wake flow and horizontal tail buffet are described. It is shown that the dynamic pressure data on the tail surface at the specified flight conditions serve to determine the local dynamic loads. They also influence the control performance of the aircraft under maneuver conditions where buffet is encountered. The data presented demonstrate a number of contributing factors that affect the tail dynamic pressures in the transonic regime.

Hwang, C.↗

NASA NTRS1978

Apparatus for determining surface tension

System for studying capillary action uses pressure transducer and chart recorder instead of manometer. Apparatus enables measurements to be made under controlled atmospheres. It also may be remotely operated. These features are particularly useful when dealing with noxious liquids and for study of surface tension under high-pressure conditions that require use of all-metal apparatus.

Razouk, R. E.↗

NASA NTRSDate not supplied

Design of a GRCop-42 Regeneratively Cooled Thrust Chamber Assembly and Feed System

An additively manufactured thrust chamber assembly was designed and printed using laser powder bed fusion. The NASA-developed material GRCop-42 was used for its high strength and high temperature characteristics. The combustion chamber and nozzle utilize regenerative coolant channels to achieve long duration hot fire tests. Design for additive manufacturing was used to create geometries not easily attainable with traditional machining. The thrust chamber assembly was post processed to ensure print quality and verify for testing. A feed system – encompassing all propellant lines, valves, sensors, and tanks – was designed and built to enable steady state, pressure regulated testing. Pressure transducers, thermocouples, and load cells were placed to enable measurements of the propellant properties and engine performance. Venturis and orifices were used throughout the system to control the flow rates to the thrust chamber. Aside from providing propellants, the feed system provides pressure using nitrogen, and purges the lines during shutdown. The design process for this testing platform is described. A testing campaign on the hardware will be conducted by the Akronauts Rocket Design Team at the University of Akron in Spring 2023.

Dillon M. Petty↗

NASA NTRSDate not supplied

Design of a GRCop-42 Regeneratively Cooled Thrust Chamber Assembly and Feed System

Dillon M. Petty↗

NASA NTRS1994

Analysis of F/A-18 Tail Buffet Data Acquired in the 80- by 120-Foot Wind Tunnel

Tail buffet studies were conducted on a full-scale, production, F/A-18 fighter aircraft in the 80- by 120-Foot Wind Tunnel of the National Full-Scale Aerodynamic Complex at NASA Ames Research Center at Moffett Field, California. Tail buffet data were acquired over an angle-of-attack range of +20 deg to +40 deg, a side-slip range of -16 deg to + 16 deg, and at wind speeds up to 100 knots. The maximum speed corresponds to a Reynolds number of l2.3 x l0(exp 6) based on mean aerodynamic chord and a Mach number of 0. 15. The port, vertical tail fin was instrumented with ninety-six surface-pressure transducers, arranged in six by eight arrays, on each side of the fin. ne aircraft was also equipped with a removable Leading-Edge Extension (LEX) fence whose purpose is to reduce tail-buffet loads. Current analysis methods for the unsteady aerodynamic pressures and loads are described. Only results for the zero side-slip condition are to be presented, both with and without the LEX fence. Results of the time-averaged, power-spectral analysis are presented for the tail fin bending moments which are derived from the integrated pressure field. Local wave velocities on the tail surfaces are calculated from pressure correlations. It was found that the LEX fence significantly reduces the magnitude of the root-mean-square pressures and bending moments. Scaling and repeatability issues are addressed by comparing the present full scale results for pressures at the 60%-span and 45%-chord location with previous full-scale F/A-18 tail-buffet test in the 80- by 120- Foot Wind Tunnel, and with several small-scale tests. The comparisons show that the tail buffet frequency scales very well with tail chord and free-stream velocity, and that there is good agreement with the previous full-scale test. Root-mean-square pressures and power spectra do not scale as well as the frequency results. Addition of a LEX fence caused tail-buffet loads to be reduced at all model scales.

James, Kevin D.↗