Search

The dynamics of unstart in a floor-mounted inlet-isolator model in a Mach 5 flow are investigated experimentally using particle image velocimetry and fast-response wall pressure measurements. The inlet compression is obtained with a 6-deg ramp and the isolator is a rectangular straight duct that is 25.4 mm high by 50.8 mm wide by 242.3 mm long. Unstart is initiated from the scram jet mode (fully supersonic in the isolator) by deflecting a motorized flap at the downstream end of the isolator. With the flap fully down, the particle image velocimetry data of the started flow capture the characteristics of the isolator boundary layers and the initial inlet reflected shock system. During unstart, the unstart shock system propagates upstream through the inlet-isolator. The particle image velocimetry data reveal a complex, three-dimensional flow structure that is strongly dependent on viscous mechanisms. Particularly, the unstart shock system propagates upstream and induces significant boundary-layer separation. Side-view particle image velocimetry data show that the locations of strongest separation during unstart correlate with the impingement locations of the initial inlet shock as it reflects down the isolator. For example, in the middle of unstart, the unstart shock system is associated with massive separation of the ceiling boundary layer that begins where the first inlet shock reflection impinges on the ceiling. The observation that separation increases at the inlet shock reflection impingement locations is likely due to the fact that the boundary layers in these locations are subject to larger adverse pressure gradients, thus making them more susceptible to separation. During the unstart process, large regions of separated flow form near the floor and ceiling with reverse flow velocities up to about 0.4[U.sub.[infinity]]. These regions of separated, subsonic flow appear to extend to the isolator exit, creating a path by which the isolator exit boundary condition can be communicated upstream. Plan-view particle image velocimetry data show the unstart process begins with separation of the isolator sidewall boundary layers. Overall, the unstart flow structure is highly three-dimensional. DOI: 10.2514/1.J050037

Experimental pitot pressure measurements in impulse facilities have typically had large-scale harmonic fluctuations associated with them. A combined experimental and numerical approach is used to investigate if a Helmholtz resonance is created from the shrouding that protects the pressure sensor from particle impact, rather than from disturbances being present in the freestream. To verify this experimentally, hydraulic oil was used to change the sound speed in the cavity. Numerical calculations of the pitot probe were used to show that both the steady and the expected transient inflow cause a Helmholtz resonance of similar period. However, a 5 % isothermal level of freestream noise was required to match the experimental fluctuation levels. Viscous effects were also shown to be significant during the initial transient response of the pressure measurement. A new pitot probe design was successfully tested and shown to be able to reduce the magnitude and period of the fluctuations. DOI: 10.2514/1.42543

The effects of placing a slot through a two-dimensional Wortmann FX 63-137 airfoil in ground effect were examined using computational fluid dynamics. The geometric shape of the slot was varied in three different ways: the width of the slot (w/c = 0.02, 0.04, and 0.06), the angle of the slot with respect to the airfoil's chord line (d = 20, 30, and 40), and the position of the slot along the chord line, (x/c = O. 15, 0.20, and 0.25). In addition, the airfoil was tested at five different angles of attack: -3, 0, 5, 10, and 15 deg. The commercially available software Gambit 2.3.16 was used to create the computational grids. FLUENT 6.2.16 with the renormalized group k-e turbulence model was then used to simulate the flow. Pitch stability of the slotted airfoil was examined and results indicated that increasing the angle of attack of the slotted airfoil while in ground effect had a reduced increase in lift when compared with the lift generated by the baseline Wortmann FX 63-137 airfoil. Results also showed that the slot could be used to reduce center-of-pressure movement along the chord of the airfoil for the range of angles of attack investigated, thus improving the overall pitch stability of the airfoil. The slot geometry that produced a minimal center-of-pressure fluctuation was located at 20 % of the chord length from the leading edge with a width of 2 % and an angle of 20 deg between the slot and a line normal to the chord line. DOI: 10.2514/1.31246

Wall pressure and wall pressure fluctuations in turbulent flows are of interest in many engineering applications. This paper focuses on the design and fabrication of aero-micro-electromechanical system surface pressure sensor arrays and their application to wall pressure measurements on a wall-mounted cylinder. The sensor arrays have been developed to be highly sensitive and mounted flush with the measurement surface. Thus, dynamic properties of the sensors are not limited by tubing, and an accurate measurement of the wall pressure and its fluctuations becomes possible. The arrays introduced herein consist of a number of individual sensors, which feature a maximum sensitivity of 12 [micro]V/(VPa) and a pressure resolution of up to 0.5 Pa. Employing six arrays (three different types) consisting of up to 13 individual pressure sensors, wall pressure measurements have been conducted at the University of Rostock at a Reynolds number of 200,000. The results prove that the developed sensor arrays are a powerful measurement tool for experimental fluid mechanics. The wall pressure measurement data obtained are in good agreement with the results of the laser Doppler anemometry measurements and large-eddy simulations, which are also presented herein.

The present investigation focuses on assessing the predictive capabilities of state-of-the-art computational fluid dynamics for a generic fan-in-wing configuration by comparison with experimental data. The objective is to reproduce a short take off and landing or a transition-flight situation without ground effect. A rotating fan is placed in the wing plane, inside of the wing's rear part and close to the root section. The obtained experimental data include force measurements, surface pressure measurements, flowfield mapping using particle image velocimetry and wool-tufts visualization. A structured mesh of the entire configuration with minimum geometrical simplifications is used to perform unsteady Reynolds-averaged Navier-Stokes computations. The area surrounding the rotor blades is set up as sliding mesh to simulate the rotation. Because of the abrupt deflection of the flow by the fan, an unsteady recirculation area is generated above the rotor blades resulting in a highly distorted inflow. The blocking of the freestream by the jet exiting the fan's nozzle creates a back pressure affecting the internal aerodynamics. The jet rolls up in a counter-rotating vortex pair with considerable impact on the wing's aerodynamic performance. Time-averaged unsteady results over one fan revolution show a good agreement with the experimental data. The major turbulence phenomena are well predicted by the simulation.

The primary difficulty with obtaining high-fidelity pressure measurements is acoustical distortion within the various components of the pneumatic system. The method of Berg and Tijdeman is the de facto standard for the aerospace industry for calculating the frequency response of pressure-sensing systems. However, the model is complex, difficult to program, and inherently linked to the frequency domain. For many applications, a less complex time-domain model is desired, and considerable work has previously been performed on this topic. Unfortunately, existing low-order response models are not particularly accurate. These models are derived using simplifying assumptions that lose many of the salient features of the flow physics. This paper derives a second-order model directly from the Berg and Tijdeman solution and the model accurately describes the physical characteristics of the dominant response harmonic. The model is modified for application to both laminar and turbulent flow conditions. Frequency- and step-response comparisons are presented to demonstrate that the derived second-order model has superior response fidelity. Accuracies of model predictions are validated by a series of laboratory step-response experiments.

An analytical model to predict the primary stagnation pressure that starts an annular injection supersonic ejector is presented. If the length of the mixing chamber is longer than the critical length, the starting pressure increases proportionally to the mixing chamber length, which differs from conventional ejectors using central injection. To describe the dependency of the starting pressure on the mixing chamber length, we assume that the ejector starts when the exiting supersonic primary flow reaches the second throat. In the present model, we use a subsonic mixing model to calculate the secondary flow pressure in the mixing chamber. Applying the obtained pressure as a back pressure condition, the distance that the supersonic primary flow develops is calculated. Comparing the distance to the mixing chamber length, we derive a minimum pressure requirement diagram that accurately predicts the starting pressure and hysteresis for a given geometry.

Pressure-sensitive paint measurement can obtain a much more detailed surface pressure distribution than can be obtained using conventional pressure taps. However, the pressure-sensitive paint is sensitive not only to pressure but also to temperature, and where high accuracy is required, it is essential to compensate for this temperature dependency. This paper discusses data processing methods for pressure-sensitive paint measurement in transonic industrial wind tunnel testing, and proposes three methods to compensate for temperature dependency of the pressure-sensitive paint: an in situ method, an a priori method, and a hybrid of a priori and in situ methods. The pressure distributions from the pressure-sensitive paint data obtained by these proposed methods are compared with pressure tap data measured by conventional pressure transducers, and it is confirmed that the proposed methods are effective in compensating the temperature dependency of pressure-sensitive paint and improve the accuracy of the obtained data. It is also found that the hybrid of a priori and in situ methods is widely applicable to the industrial wind tunnel testing even if the pressure range of the pressure tap data is limited.

This paper presents and analyzes unsteady surface pressures measured on a model rotor blade as the blade experienced near-parallel blade-vortex interaction with a twin-vortex system. To provide a basis for analysis, the vortex system was characterized by hot-wire measurements made in the interaction plane but in the absence of the rotor. The unsteady-pressure response resulting from a single-vortex interaction is then presented to provide a frame of reference for the twin-vortex results. A series of twin-vortex interaction cases are then presented and analyzed. It is shown that the unsteady blade pressures and forces are very sensitive to the inclination angle and separation distance of the vortex pair. When the vortex cores lie almost parallel to the blade chord, the interaction is characterized by a two-stage response associated with the sequential passage of the two cores. Conversely, when the cores lie in a plane that is almost perpendicular to the blade chord, the response is similar to that of a single-vortex interaction. In all cases, the normal force response is consistent with the distribution of vertical velocity in the flowfield of the vortex system. The pitching moment response, on the other hand, depends on the localized suction associated with the vortex cores as they traverse the blade chord.

A methodology to correct for temperature effects on pressure-sensitive paint data in a blowdown wind tunnel has been suggested. PSP measurements using Optrod-B1 paint were made on an aircraft model instrumented with thermocouples and conventional pressure taps. The tests were made at freestream Math numbers of 0.6 and 0.8 and at model incidences of 6 and 10 deg. The temperature correction is based on a mean model temperature concept for a given blowdown and the corrected PSP data are validated using pressure port data. It has been found that the accuracy of temperature-corrected PSP results is about as good as for those obtained at transonic speeds in other tunnels elsewhere.

An empirical model of the surface pressure spectrum beneath a two-dimensional, zero-pressure-gradient boundary layer is presented that is based on the experimental surface pressure spectra measured by seven research groups. The measurements cover a large range of Reynolds number, 1.4 x [10.sup.3]

A two-step calibration and measurement process for seven-hole pressure probes was devised to permit simple and accurate real-time processing. The technique adds an extra step to the usual least-squares calibration procedures such as Gallington's method involving the interpolation of error loop-up tables. A measurement system using the two-step calibration method was developed. Measurements were made in pipe and wind tunnel flows and a detailed uncertainty analysis was conducted to verify the accuracy of the new method.

Miniature multihole pressure probes and a unique neural network-based calibration algorithm for them were developed. Seven-hole probes of tip diameters as small as 0.035 in. were produced with high tip surface quality. The miniature probes were calibrated and tested in a wind tunnel. A backpropagation-based neural network calibration algorithm was designed for these probes. In the feedforward mode, the algorithm provided computational speeds with an order of magnitude higher than those commonly attained by similar accuracy interpolation algorithms.

A flowfield velocity and pressure measurement probe with an almost omnidirectional velocity measurement features is developed for three-velocity-component and pressure measurements that centers on the methods used in probe fabrication, calibration, and frequency-response research. The probe tip is a sphere with 18- pressure-port spherical head evenly distributed over the surface. The design and construction of the probe and its calibration for use in compressible flows are discussed.

The luminescence of a pressure sensitive paint formulation consisting of a luminescent polypyridine ruthenium(II) dye dispersed in polydimethysiloxane binder exhibits temperature dependence effects. The intrinsic temperature dependence of the nonradiative decay rate of the luminescent probe molecule effects the temperature dependence of emission intensity and decay lifetime. The effect occurs under deoxygenated conditions. Temperature dependence of oxygen diffusivity in the polymer matrix effects the temperature dependence of the luminescence emission variables at 1-atm air pressure.

An analytic technique uses mean pressure measurements to predict fluctuating loads created by shock-wave turbulent boundary-layer interactions. It assumes that separation shock unsteadiness cause fluctuating loads and that the probability density function of the instantaneous separation shock location is Gaussian. Based on the knowledge of the location of upstream influence and separation, the technique predicts the loads with a 10% accuracy. It relates the measured mean wall pressure, the wall pressure root mean square and the intermittency. Variations of the method are discussed.

Two three-dimensional compressible Navier-Stokes solvers accurately predict the low mach-number performance of an untwisted rectangular wing and a subsonic transport wing-fuselage configuration. The computed pressure distributions are almost identical and agree well with the experimental data. The predicted pressure distributions are also extremely close to those predicted by an incompressible Navier-Stokes solver. The three flow solvers predict almost similar boundary profiles and the computational time is case dependent. The compressible solvers employ a finite volume formulation.

Three-dimensional swept-shock-wave/boundary-layer interactions produce highly fluctuating pressure loads on portions of high-speed aircraft. The regions close to the foot of the fin that generates the swept shock wave experience the highest loads. Low-frequency components dominate the fluctuations that are caused by random motion of the primary attachment line. Motion of the separation shock causes an rms pressure peak which decreases with increasing sweepback of the separation line. The causes of the fluctuations are discussed.

The numerical predictions of forebody, leading-edge extension, inlet lip and duct surface pressures in F/A-18A inlet flow compare well with the flight data. A significant degree of flow unsteadiness exists at the engine face. The inlet flow distortions and total pressure recovery at the engine face are at the upper and lower range of test data. The calculated total pressure contours are within the excursion range of the minimum total pressure indicated by the flight data.

The pressure sensitive paint (PSP) technique efficiently computes pressure on a 4.7% scale F-15E model during a high-speed wind tunnel test. Measured pressure distributions are consistent with the geometric characteristics of the model. In situ paint calibration determines absolute accuracy and quantitative utility. The test procedure of PSP technique is discussed.

A novel method developed for deep cavities suppresses cavity pressure oscillations. Grazing flow over deep cavities cause acoustic resonances. Two cases for deep cavity, involving the single vortex, fundamental frequency excitation and the double vortex, first odd harmonic excitation, are analyzed. Measurement of mode shapes permits the optimization of control strategy for the two resonant conditions.

The unsteady pressures control the momentum balance for both viscous and nonviscous fluid. The momentum generated in a compressible atmosphere by an impulsive point force is distributed between a moving fluid sphere and a front expanding with the velocity of sound. The net momentum is zero in a transverse section of an incompressible atmosphere. A comparison between the viscous boundary layer and the actuator surface of the atmosphere shows that the power taken by actuator is a fraction of the work performed by the system from outside.

Pressure distribution at an aircraft tire-wheel interface can be studied by an experimental method using piezoelectric film. The method gives accurate inflation-pressure values at discrete points by measuring voltage changes due to pressure but it is necessary to know the in-plane strain field at gauge location. Another method, the hybrid method, using the finite element technique gives continuous pressure distribution values. Analysis of the results obtained using the two techniques is given.

The pressure-sensitive paint method, which uses the oxygen sensitivity of specific photoluminescent materials as paint, is versatile in mapping the pressure field over aerodynamic surfaces. Among other features, quantitative pressure measurements over various points on the model surface can be obtained through this methodology.

Surface pressure measurements on pitching airfoils up to angles of incidence greater than stall angles yield spatially well-resolved pressure distributions that agree well with experimental data. The obtained pressure magnitudes are greater than previous experiments at low Reynolds numbers due to fast initiation of laminar-turbulence in the flow. The experiment is conducted in a water channel which yields Reynolds numbers of 10 to the power 5 at low velocities of about 30 meters per second.

Pressure sensitive luminescent coatings are used in a novel surface pressure measurement method that can be used in aerodynamic applications. The effect of accurate measurements of input quantities on the accuracy of pressure for these pressure-sensitive paints is studied. These paints obey the Stern-Volmer relation, and the ratio technique functions at optimum levels at moderate pressures and moderate-to-high Mach numbers.

The Navier-Stokes code is used with both the Baldwin-Lomax model and a mixing length turbulence model to investigate pressure coefficients on a typical O-grid around a blunt trailing-edge. The O-grid, instead of the more complex C-H grid, is used as the computational domain. Both models give very similar results that are at least as good as those obtained using one-equation turbulence models.

The unsteady nature of propeller-wing interaction, studied using flow visualization of steady and unsteady wing surface pressure measurements, involves the measurement of wing pressures over a wing fitted with a chordwise array of 16 microphones in a spanwise direction through the propeller wake. The propeller tip vortex is marked with smoke, enabling a qualitative study of the physical nature of the propeller tip vortex motion over the wing, and aiding the interpretation of unsteady pressure measurements.

The estimates of entropy production and pressure variation due to turbulent mixing were computed by using a quasi-one-dimensional model. The model has been previously used to prove that large losses in pressure occur when compressible mixing takes place with a large velocity difference. Most applications employ the method of mixing air and hydrogen as in scramjet engines. This will inevitably lead to a large delta U formulation over the area of operation of the device. The high delta U induces faster mixing is needed for combustion.

A conical flow asymmetry exists for extremely slender cones, where the asymmetric flow separation begins to develop at an angle of attack much below 30 degrees. The study is based on experimental results derived from analyzing slender bodies of revolution at high angles of attack. When the angles of attack is 30 degrees or more, the asymmetry is not conical in laminar flow condition, but displays characteristics similar to that taking place on cylindrical afterbodies. An asymmetric vortex pair can only be developed from symmetric cross flow separation if symmetric cross flow attachment is restrained.

The unsteady three dimensional Euler equations are solved to determine the unsteady blade pressures caused by the operation of the propfan at an angle to the direction of the mean flow. The eight-bladed SR7L propfan operating at takeoff conditions is considered with the inflow angles as 6.3, 8.3 and 11.3 degrees. The distortion is caused by viscous separation effect, and the extent of distortion is directly proportional to the inflow angle.

Pressure and velocity measurements and flow visualization are used in the study of the behavior of a planar jet impinging on a V-shaped plate. At a distance L from the jet exit downstream, the plate with a symmetrically variable opening angle is located. The existence of large scaled coherent structures in the shear layers in the jet and the exhibition of a prominent periodic flapping motion with a preferred frequency in the flowfield in the jet column are related to the presence of the plate.

The pressure distribution on a prop-fan blade was computed using a Total Variation Diminishing algorithm of the Euler code. The aim of the study was to show correspondence between measured and experimental data at various working conditions. Wind tunnel tests were conducted on an SR-7L prop-fan blade driven by Twin Turbomeca gas turbine engines with a common gearbox. Results show that the computed pressure distribution and power coefficient agree with experimental data for Mach numbers of 0.5 and O.78. The Euler code used has been demonstrated of capably designing the prop-fan blade.

A new procedure for correcting the error produced during the measurement of the pressure in miniature multihole pressure probes has been developed. The error is caused by the lag in the response of the probe tubing system. The frequency response of the probe deteriorates as the size of the pressure probe is reduced to lessen flow disturbance. The novel scheme is not computationally intensive and can be implemented in real time. Howeve, its usefulness is confined only to damped or overdamped tubing systems.

A dynamic model for pressure sensing configurations in high Knudsen number hypersonic flows was developed. The model can be used for both continuous and rarefied flow conditions, and takes into account longitudinal temperature gradients in the sensor configuration. The applicable flow conditions for the model were examined in a series of laboratory tests. Model comparisons were excellent for Knudsen numbers up to 0.65. For values greater than 0.65, free molecule dynamics dominate and the model is no longer applicable.

A study of vortex breakdown over delta wings shows that the vortex breakdown position can be actively controlled by pressure variations initiated by helical mode instability of the breakdown. Flow visualization analysis for a water channel's variable sweep angle reveals a normalized time constant of tau-U(sub infinity)/c = 2-7.

The claim that total pressure always decreases in frictional flow is wrong. Stresses can increase local total pressure and this is shown by studying a total pressure transport equation. The stresses cause redistribution of mechanical energy and produce higher total pressure, directly proportional to viscosity. Two classical flow cases that have exact solutions are discussed.

The effects of streamwise pressure gradient on the growth rate of a supersonic mixing layer are investigated through experimental procedures. Tests are carried out in a shock tunnel wherein two types of nozzles which generate the mixing layers are independently subjected to simulated supersonic flow. The necessary measurements are then obtained by means of Schlieren photography. Results reveal that the growth rate of the mixing layers is enhanced by the presence of the streamwise pressure gradient.

A unique vortex generator that simulates the trailing vortex produced by a helicopter main rotor in forward flight was used in a study of main rotor/tail rotor blade vortex interaction. Velocity measurements were performed using hot wire anemometry and particle image velocimetry, which showed the interacting vortices to be three dimensional and to contain a significant core axial flow. Surface pressure measurements during vortex collision with an instrumented, stationary blade were obtained at a variety of incidence settings. At the leading edge, a suction peak occurred where the core flow was directed away from the surface and, similarly, a pressure peak occurred where the core flow was directed toward the surface. This behavior can be explained in terms of the impulsively blocked vortex core axial flow. Also, the leading edge suction and pressure peaks were amplified or attenuated by changing the incidence of the blade. However, the changes in surface pressure distribution were such that the peak normal force was approximately the same for each incidence setting.

An analysis of axial mean velocity and static pressure distributions in the near field of a circular jet reveals small deceleration of centerline mean velocity and corresponding acceleration of static pressure inside the potential core near the exit. The streamwise mean velocity profile has a saddle-shape behavior which is limited to the potential core. The mean static pressure profile, corresponding to the saddle-shape mean velocity profile, shows substantial drop.

A theory of the unsteady motion of the cascades of vortex sheets suggests that the pressure in the far wake of rotor blades is equal to the ambient pressure. The consideration of the unsteady motion of the vortex sheets creates a similarity between the results of the unsteady theory and the impulse-momentum theory. An infinite cascade of discrete vortex rows forms the base for examining the unsteady theory. Results from the calculation of steady pressure and unsteady pressure using Bernoulli's and Kelvin's equations agree with theoretical ones for 10 vortices.

A technique for conjuring qualitative gas temperature distributions surrounding hypersonic automobiles was proposed. The vibrational and rotational mechanisms of molecules change in accordance with the changes in the gas temperature. The excitability in gas molecules differ, and electric discharge from the flow field of gas temperature variables produces greater excitability in gas molecules. The gas temperature distribution can be recorded on film.