Model Surface Pressure Research Articles

Viscous drag reduction using riblets on NACA 0012 airfoil to moderate incidence

Results of viscous drag reduction using riblets from 3M on a NACA 0012 airfoil model up to moderate angles of attack are presented. Measurements made consisted of model surface pressure distributions, mean velocity and streamwise turbulence intensity profiles in the boundary layer (just ahead of the trailing edge), and total airfoil drag for two riblet heights of 0.152 and 0.076 mm. Results show significantly higher skin friction drag reduction with incidence compared to plat plate flows; the reduction was as high as 16% at ? = 6 deg. Results of mean velocity profiles show that a larger contribution to drag reduction results from the suction side of the airfoil, indicating increased effectiveness of riblets in adverse pressure gradients. Examination of turbulence intensity profiles in the wall region indicates an appreciable reduction in the presence of riblets; correspondingly, the spectra show reduced energy levels at low frequencies.

Scramjet nozzle experiment with hypersonic external flow

An experimental study of scramjet asymmetric nozzle flowfields is conducted to investigate the effects on nozzle performance of interactions between engine exhaust and hypersonic external flow. The test model consists of a flat-plate ramp and a short cowl. Tests are performed in a hypersonic wind tunnel (Mx = 7.1) whose freestream constitutes the external flow around an aerospace plane. Room temperature air is used to simulate engine exhaust flow. Model surface pressure and pitot pressure distributions in the flowfields are measured, from which nozzle performance is calculated. Flow visualizations are performed using several techniques. It is found that in underexpansio n conditions, the external flow affects the model surface pressure only from the ramp sides because the intercepting shock (barrel shock) does not impinge upon the ramp surface. It is shown that the external flow produces an effect which suppresses boundary-layer separation on the ramp surface. When using a long side fence (SF), span wise expansion of the exhaust flow is restrained, thus, a slight gain of thrust can be obtained relative to that using a shorter side fence. The predicted nozzle performance by the twodimensional method of characteristi cs with boundary-layer correction agrees well with the long side fence experimental results.

Ring evolution in general circulation models from path analysis

The thesis explored here is the use of path data for comparing primitive equation models with observations. Such comparisons naturally focus on the inversion of path data for fundamental kinematic characteristics and subsequent descriptive interpretations of both models and observations. The analysis here focuses on uncertainties that may affect this interpretation. These include random position errors and numerical noise introduced by the nonlinear inversion ansatz as well as dynamical processes that introduce both low‐ and high‐frequency variability in the kinematic fields. After each of these effects is assessed, the inversion ansatz is applied to paths generated from an eddy‐resolving general circulation model of the Gulf of Mexico. Model Loop Current ring centers determined from different paths are found to be consistent and in good agreement with observations in the abyssal gulf. In addition, these centers are compared with the model surface and interface pressure anomaly centers. The simulated paths showed ejection of mass from the Loop Current ring as it interacted with the western slope. Observed and simulated ring kinematics inferred from appropriate paths are also made and compared with observations for the slope region. The paper concludes with a discussion of the potential for using simulated and observed path data for extracting flow physics and quantifying comparisons between models and observations.

Pressure measurement in a blow-down wind tunnel

An automatic system for steady pressure measurement in a small hand-controlled supersonic wind tunnel is described. The ratio of model surface pressure to tunnel blowing pressure was measured using a Scanivalve controlled by a small digital computer. The software contained provision for the rejection of samples which were not sufficiently steady, and a transducer calibration routine was included to enable a variety of transducers to be used with the system. Typically, 18 pressure tappings could be sampled in a 50 s run (including start-up time) with an error of ±2% in the pressure ratio. The system should be applicable to any situation where run times are limited and difficulty is experienced in maintaining precise control over blowing pressure.

Aerosound from corner flow and flap flow

Noise generation at the edge of a wing flap is analyzed. The phenomenon as a single vortex moving around a corner in an incompressible, potential flow is modelled. Vortex image retarding effects are proposed as an explanation for small Strouhal numbers. The model surface pressures, sound pressures (using Curle's theory), and Mach number dependencies agree with wind tunnel experiments. A double pressure peak is found in the model (credited to image action) which is qualitatively similar to measured sound correlations. Incompressible flow aerosound calculations are discussed. The effects of a series of vortices moving in the same idealized potential flow are also studied. The vortices are assumed to be statistically independent so their intensities can be added. The frequency of appearance of the vortices are determined from measurements. Diffraction effects caused by the presence of the wing near the dipole sound radiators on the flap surfaces are included.

Aerosound from corner flow with application to the flap noise problem

Recent correlation measurements have indicated that the major jetplane airframe noise source is located at the outboard edge of deployed flaps. The surface noise generated there is apparently due to separation and disruption of the spanwise flow at the underside edge of the flap. We model this using incompressible potential flow around a corner, with a series of vortices moving about the corner, imbedded in the flow. The model surface pressures, sound pressures (using Curle's theory) and Mach number dependencies agree well with wind tunnel experiments. A double pressure peak is found in the model (credited to image action) which is qualitatively similar to measured sound correlations. We discuss incompressible‐flow aerosound calculations. Corrections are applied for the diffraction effect of the wing on the dipole surface sound (the effect enhances the forward‐radiated field). We suggest that a possible explanation of the typically small Strouhal numbers seen in these experiments, is the retarding effect of the vortex image, on the motion of the vortex in the flow. [Work supported by NASA‐AMES Research Center.]

Trailing-Edge Flows at High Reynolds Number

An investigation of trailing-edge flows at high Reynolds number and subsonic Mach numbers is presented. Symmetric and asymmetric trailing-edge flows are studied, each flow having pressure gradient regions upstream of the trailing edge similar to an airfoil. Measurements include model surface pressures, mean velocity, turbulent shear stress, and turbulent kinetic energy profiles in the trailing-edge and near-wake regions. Comparisons of the symmetric data with numerical solutions of boundary layer as well as Navier-Stokes equations employing two different turbulence models show increasing effects on viscous interactions as the Mach number increases. Both turbulence models yielded solutions of the mean flow of comparable quality. The experimental results of the asymmetric case are discussed.

High-speed monorail rocket sleds for aerodynamic testing at high Reynolds numbers.

The use of a monorail rocket-sled system is detailed and recommended as a method of obtaining aerodynamic data in a regime of high Mach and Reynolds numbers. Sled designs and the data collection system are discussed in detail. The system has been used successfully to acquire pressure data on cone-cylinders. These data are compared with data from free-flight rockets, wind-tunnel tests, and computer calculations. The sled-testing technique appears to be the most promising method of simulating flight conditions at low altitudes at Mach numbers up to M = 8. A principal advantage is the high degree of accuracy possible in determining actual test conditions, which include Mach number, static pressure, and static temperature. Exploratory tests indicate that sleds may be used to collect other types of experimental data such as heat-transfer and boundary-layer transition. To date, excellent data have been obtained at 7430 fps (M = 6.53), which to our knowledge is the highest velocity ever attained by a rocket sled. At this speed, and the Holloman Air Force Base track altitude, the Reynolds number was 42 X 106/ft, and the dynamic pressure was 55,000 psf. Nomenclature CD, CL = drag and lift coefficients, respectively, based on sled cross-sectional area M = Mach number Pm = measured model surface pressure POO = freestream ambient pressure Re = freestream Reynolds number To, Tw, Tm = total, wall, and freestream temperatures, respectively U = freestream velocity, fps

Experimental investigation of supersonic laminar, two-dimensional boundary-layer separation in a compression corner with and without cooling.

An experimental investigation of the boundary layer separation associated with a compression corner was conducted in the GALCIT Mach 6 wind tunnel, and a supplementary study was performed in the JPL supersonic wind tunnel. Special emphasis was placed on the development of a wind tunnel model which approximated true two-dimensional flow, and which could be run in either a highly cooled or an adiabatic configuration. The basic measurements consist of the model surface pressure and temperature, and Pitot surveys of the boundary layer. The surface pressure distributions for the adiabatic wall configurations are compared with the theory of Lees and Reeves (modified by Klineberg and Lees). The surface pressure distribution for the cold wall was compared with the adiabatic configuration for a laminar interaction, and the dependence on Reynolds number for both laminar and transitional interactions are observed. The free interaction similarity suggested by Chapman is empirically tested and found to be a good approximation for the adiabatic configuration, but it fails to correlate the cooled with the adiabatic case. The scaling suggested by Curle was tested and found to eliminate this deficiency.

Model Surface Pressure Instrumentation for a Hypersonic Shock Tunnel

The design and construction of piezoelectric pressure transducers to meet model surface pressure instrumentation requirements for the Grumman Research hypersonic shock tunnel are described. An accurate method for calibrating the gauges is given. The gauge response to calibration pressures and to pressures measured in the tunnel is analyzed to account for the relatively short time constant of this system.