Freestream Dynamic Pressure Research Articles

Planform Variations and Aerodynamic Efficiency of Movable Tip Strakes

b = span, mm c = chord, mm D = wing drag, lb CL = wing lift coefficient CD = wing drag coefficient d = angle relative to wing chord plane, deg L = wing lift, lb q = 1 2 1V 1 freestream dynamic pressure, kPa Re = 1V1c= 1 Reynolds number based on chord S = reference area, mm V1 = freestream velocity, m=s x, y, z = aerodynamic axes = angle of attack, deg = leading-edge sweep, deg = absolute viscosity, N s=m = air density, kg=m

Time-Series and Time-Averaged Characteristics of Subsonic to Supersonic Base Flows

Physics of cylindrical base flows ranging from subsonic to supersonic speeds at zero angle of attack are computationally investigated. Time-series and time-averaged investigations of base flows show distinctive characteristics at subsonic, transonic, and supersonic regimes. Normalized time-averaged base pressure decreases proportionally with respect to increasing freestream dynamic pressure in the subsonic regime of M ∞ 1.5. Normalized base pressure fluctuations sharply increase at transonic speeds, whereas they decrease with increasing freestream Mach number at subsonic and supersonic speeds. Appearance of unsteady local shock waves change the characteristics of base pressure distinctively at the transonic speeds. Spectra of the base pressure show one clear peak at subsonic speeds (related to the shear layer dynamics), two clear peaks at transonic speeds (related to the shear layer dynamics and its subharmonic), and three major peaks at supersonic speeds (related to the shear layer dynamics, its subharmonic, and an additional mechanism). Instability of the free shear layers has dominant influence on the overall base flowfield over a wide range of Mach numbers ranging from subsonic to supersonic speeds. However, at supersonic speeds, an additional mechanism of instability within the recirculating region is possibly at work and has dominant influence on the flowfield. The dominant mechanisms significantly cause the strong Mach number dependence of the high-pressure region which is strongly related to the base pressure. The spectrum analysis suggests that the substantial base pressure fluctuations are caused by the pulsing of the flow inside the recirculating region.

Separation Control on a Trailing-Edge Flap Using Air Jet Vortex Generators

Nomenclature A = area CD = drag coefficient CL = lift coefficient Cp = pressure coefficient CV = jet velocity ratio Cμ = blowing coefficient (jet momentum ratio) c = chord DLG = flap deflection, lap and gap f = jet orifice diameter L = lift M = Mach number n = number of jets p = static pressure p0 = total pressure q = freestream dynamic pressure S = wing reference area V = speed x/c = nondimensional distance measured in a chordwise sense α = angle of attack, or jet skew angle β = jet pitch angle γ = ratio of specific heats = flap deflection λ = jet orifice spacing ρ = density

Accelerating CFD-Based Aeroelastic Analysis Using Distributed Processing

Distributed processing is applied to CFD-based aeroelastic analysis to accelerate the completion of flutter boundary predictions. The process for identifying an aeroelastic stability boundary is divided into several independent simulations that are distributed to multiple computers. The solutions are completed in parallel. This parallelism is used to accelerate the prediction of the instability boundary. This process is also applied to the generation of training data for the system identification of the unsteady aerodynamic response to motion and force on a modal structural model. The equipment used in this study is common, off-the-shelf computing hardware. This method is application ready with no required alterations to the CFD solver. Results address the extent to which this methodology is applicable to aerospace applications. Nomenclature f a = generalized aerodynamic force vector N = number of data points in training set Np = number of data points in a parallel training set na = number of previous aerodynamic forces nb = number of previous structural displacements np = number of Central Processing Units nr = number of roots or mode shapes q = free stream dynamic pressure q = generalized displacement vector ρ = free stream density

Three-dimensional coherent structure in a separated and reattaching flow over a backward-facing step

An experimental study was carried out to elu- cidate the large-scale vortical structure in a separated and reattaching flow over a backward-facing step. The Rey- nolds number based on the step height (H) was ReH=33,000. The large-scale vortical structure was probed by means of three-dimensional velocity measurements performed at the recirculation zone (x/H=4.0) and the reattachment zone (x/H=7.5). A 32-channel microphone array extending in the streamwise and spanwise directions was used for sensing the wall pressure fluctuations. The relationship between the flow field and the relevant spatial mode of the pressure field was determined by examining the spatial box filtering. From the relevant spatial mode of the wall pressure fluctuations, a conditional averaging technique was employed to characterize the coherent structure. In addition, the cross-correlation between velocity and wall pressure fluctuations was calculated to identify the structure and the length scale of the large-scale vortex. The cross-correlation results revealed that the large-scale hairpin vortices have a three-dimensional structure, in agreement with previous findings. The pres- ent results clearly show the growth and downstream elongation of the hairpin vortices. List of symbols H step height, m k turbulent kinetic energy, m 2 /s 2 q freestream dynamic pressure, Pa ReH Reynolds number based on U0 and H,UoH/m U0 freestream velocity, m/s Uc convection velocity, m/s X0 streamwise coordinate of the measurement origin, m xR time mean reattachment length, m

Vortex Generator Installation Drag on an Airplane Near Its Cruise Condition

A method is discussed for predicting the drag increment caused by the installation of a blade-type vortex generator (VG) on a transonic-transport airplane. The original Nash and Bradshaw magnie cation concept of roughness drag is extended to cover compressible e ows and then is applied to VG blades to estimate the VG installation drag on an airplane. Thedrag of a VG blade placed on a wing will beamplie ed dueto thegrowth of the boundary layerwith distance along thewing surface. Nash and Bradshaw showed that thedegree of magnie cation cannot be approximated simply by the ratio of local to freestream dynamic pressure ( q effect). To demonstrate the magnie cation effects, some VG installation drag analyses for transonic-transport airplane models are performed using the new magnie cation factor formula. It can be seen that the agreement between these predicted drag increments and the wind-tunnel test results is good, but the drag increment based on the q effect is seriously underestimated.

The structure of wall-pressure fluctuations in turbulent boundary layers with adverse pressure gradient and separation

Space–time correlations and frequency spectra of wall-pressure fluctuations, obtained from direct numerical simulation, are examined to reveal the effects of pressure gradient and separation on the characteristics of wall-pressure fluctuations. In the attached boundary layer subjected to adverse pressure gradient, contours of constant two-point spatial correlation of wall-pressure fluctuations are more elongated in the spanwise direction. Convection velocities of wall-pressure fluctuations as a function of spatial and temporal separations are reduced by the adverse pressure gradient. In the separated turbulent boundary layer, wall-pressure fluctuations are reduced inside the separation bubble, and enhanced downstream of the reattachment region where maximum Reynolds stresses occur. Inside the separation bubble, the frequency spectra of wall-pressure fluctuations normalized by the local maximum Reynolds shear stress correlate well compared to those normalized by free-stream dynamic pressure, indicating that local Reynolds shear stress has more direct influence on the wall-pressure spectra. Contour plots of two-point correlation of wall-pressure fluctuations are highly elongated in the spanwise direction inside the separation bubble, implying the presence of large two-dimensional roller-type structures. The convection velocity determined from the space–time correlation of wall-pressure fluctuations is as low as 0.33U0 (U0 is the maximum inlet velocity) in the separated zone, and increases downstream of reattachment.

Approximating Collisional Freestream Attenuation at Transitional Knudsen Numbers

When a body such as an instrument package is immersed within a rarefied flowfield, the inevitable result is distorsion of the freestream. This perturbation is a formidable stumbling block for code developers seeking reliable experimental data for comparison with flowfield predictors. Current codes produce a spatial map of dynamic pressure. Validation of the codes must rely on an accurate means of determining the undisturbed freestream dynamic pressure from experimental data for comparison. Althrough this problem is closed for a continuum case via Rayleigh-Pitot formula, the extreme low density inherent to an exoatmospheric plume renders this treatment invalid because the thickness of the shock layer may be many times larger than the body immersed in the flow. The basic approach taken to solve this problem was to analyze the interaction between two separate streams of molecules.

F-15 forebody vortex flow control using jet nozzle blowing

A low-speed wind-tunnel test was conducted to determine the lateral-directional control effectiveness of a forebody jet nozzle blowing on a 10% scale F-15E model. This investigation of pneumatic forebody vortex flow control was the first of its kind performed on an F-15 configuration. The test acquired sixcomponent force balance data in the NASA Langley 30- by 60-Foot Tunnel at a freestream dynamic pressure of 8 psf at an equivalent Reynolds number per foot of approximately 5.08 X 10 5. Asymmetric forebody jet blowing was effective in generating yawing moments through an angle of attack (AOA) range that extends beyond the limit of rudder control power. The optimum nozzle configurations were pointed outward, and yawing moments generated within the maneuver range of AOA (a < 40 deg) increased with increases in both blowing rate and AOA. Above 10-deg AOA, blowing rates of C M = 0.010 and greater resulted in yawing moments that approached or exceeded that developed by full rudder deflection.

Interference galloping of a circular cylinder near a plane boundary

The galloping motion of an elastic circular cylinder, with oscillations restricted to a plane normal to the incident uniform flow, when it is located in the close neighbourhood of a plane boundary is investigated theoretically by using a quasi-steady assumption; that the cylinder and the flow are both parallel to the boundary. Closed form expressions are presented to describe the cylinder behaviour: namely, the variation of the reduced (relative) amplitude with the reduced velocity, the relative mean position of the vibration with the dimensionless free stream dynamic pressure and the relative motion frequency with the reduced amplitude. The first variation and also that of the cylinder mean position with the flow velocity approach asymptotically constant values, the second one is a function only of the lift coefficients, while the last one depends on both the dimensionless free stream dynamic pressure and the lift coefficients. The mathematical model predicts that an increase in the cylinder mass or structural damping causes the galloping to occur at a higher flow velocity. Simple expressions are presented for the prediction of the reduced velocity and the cylinder motion frequency at the galloping onset. The theory indicates that the galloping instability occurs only when the gap between the cylinder and the boundary is very small.

Some readily quantifiable aspects of scramjet engine performance

Some aspects of scramjet engine performance are examined that can be treated quantitatively within the scope of a channel flow approximation including real-gas chemistry and reaction kinetics, in order to establish an accurate, traceable, maximum performance baseline for consideration of the more difficult technical issues such as mixing in the combustor, cooling, and the aerodynamics of the inlet and the nozzle. The analysis shows that if there is a lack of complete recombination in the nozzle, then there is a combustor inlet Mach number M4 that maximizes 7sp, for flight Mach numbers above about 10. The optimum M4 increases as the nozzle recombination decreases. A combustor pressure of about 1 atm is required for kinetics fast enough to ensure completion of the combustion in a length of the order of 3 ft and, with this constraint, operation at the optimum combustor inlet Mach number imposes a freestream dynamic pressure that may be unacceptably large at high flight Mach numbers. Reducing the dynamic pressure while maintaining the combustor pressure at 1 atm requires reducing M4, which, by increasing the combustor inlet temperature, degrades the specific impulse significantly from that for optimum combustor inlet Mach number. Considering all of these factors, a realistic upper limit for the vehicle mass accelerated to orbital speed from zero speed in a single-stage vehicle powered by the scramjet is about 0.25 of the initial mass. This value, which must include the structural weight of the vehicle plus the payload, would be reduced by the effects of incomplete mixing in the combustor, incomplete expansion of the exhaust nozzle, stratification of the flow in the combustor and in the nozzle, and perhaps by other effects, which the present analysis does not attempt to evaluate.

The effects of icing on the dynamic response of thick aerofoils

SummaryVibration of a flexible wing strut of 30% thickness/chord ratio occurred during icing trials of a commuter aircraft. This was caused by the formation of ice on the leading-edge of the strut producing a ‘double-horned’ ice shape. The two horns formed one above the other, having very sharp leading-edges with radii less than 0·6% of the strut chord of 0·23 m.Wind tunnel testing demonstrated that, over a limited range of incidence angle, the flow around the flexible strut was dominated by a long separation bubble that periodically ‘burst’ and re-attached. This occurred at a reduced frequency of nominally 0·06, based upon the chord of the profile and the free stream velocity. The periodic flow was also spatially highly coherent, producing relatively large unsteady forces compared with the buffet forces generated by incoherent turbulent separated flows.A high-speed film was made using smoke visualisation of the flow around a 40% rigid scale model of the strut. This was used to correlate the unsteady pressure field, measured at 18 points around the profile, with the dynamic behaviour of the separation bubble.The periodicity was found to be inherent in the flows over both the flexible and rigid struts for a limited range of incidence, occurring even when structural response was restrained to a level below that at which lock-in occurred between the flow and the response motion. The peak value of the unsteady pressure was measured on the top surface of the upper ice horn. The root mean square value of 18% of the free stream dynamic pressure was measured, with response restrained to below the lock-in threshold.

Correlating Single Phase Flow Measurements With Observations of Trailing Vortex Cavitation

The single phase and cavitating tip vortex shed by a NACA 66-209 rectangular plan-form, rounded tip hydrofoil has been studied. Single-phase measurements of instantaneous flow velocity were made by taking double-pulsed holograms of microbubbles moving in and around the vortex core. The tailored air bubble technique of Ooi and Acosta (1983) was employed to measure both the mean and fluctuating single phase vortex core static pressure. Cavitation inception was determined visually. The flow in the vortex core is highly unsteady; the r.m.s. axial velocity fluctuation can be as high as 0.2U∞. Core pressure fluctuations greater than the freestream dynamic pressure have been observed. These fluctuations are commensurate with the axial velocity unsteadiness. Cavitation inception in saturated water occurs at mean core pressures above the vapor pressure. However, the pressure fluctuations are certainly substantial enough to account for these elevated inception indices. The rapid establishment of fully developed trailing vortex cavitation when the cavitation number is reduced below inception is attributable to the small variation in mean core pressure with downstream distance. The inception index increases substantially with dissolved air content. The cause of this dependence is not presently known.

Effect of Helmholtz resonators on boundary-layer turbulence

HELMHOLTZ resonators were investigated as a passive device for the purpose of modifying a turbulent boundary layer. Resonators, imbedded in a wall, also occur inadvertently when porous walls are used to reduce boundarylayer growth or as acoustic control devices in inlet ducts and combustors. In all cases it is of interest to know the effect of the resonators on the boundary layer. Helmholtz resonators are formed by cavities that are vented by a small orifice (see Fig. 1). They respond to excitation from the shear layer at the orifice as well as from incident acoustic waves. Previous experimental work '~4 has focused on the conditions for resonance in a single resonator. In the present study, a row of 10 resonators was positioned across the span of a wind-tunnel wall. Separate studies5'6 were conducted to determine the response of the resonator row to the boundary layer and the interaction that occurred between adjacent resonators. The relevant results are that peak response was achieved at a freestream velocity of 26.5 m/s, and the frequency and amplitude of the cavity pressure oscillations were 570 Hz and 143 dB (equivalent to 80% of the freestream dynamic pressure). In addition, adjacent resonators showed a strong antiphase locking with an average coherence coefficient of 70% and a phase lag of 150 to 180 deg. The phase locking between resonators was localized, and coherence decreased rapidly with resonators that were further away. Contents Experiments were conducted in a wind tunnel with a test section of 0.71 m x 1.0 m and 2.7 m long (28 in. x 40 in. x 9 ft). Ten resonators were formed with orifices of diameter D = 1.03 cm (0.406 in.) and 0.305 cm (0.120 in.) thickness. They were spaced 3.05 cm (1.20 in) on centers occupying the central 27.5 cm (10.8 in) of the 1.0 m (40 in.) span. At the location of the resonators, with flow parameters chosen to produce a resonant condition, the boundary layer had a thickness of d = 2.8 cm (1.1 in.), friction velocity UT = 1.0 m/s, and freestream speed of 26 m/s (58 mph). The corresponding Ret> was 5500. Measurement locations were chosen at the streamwise (X) positions X/D= -3,-0.5, 0.5, 3, 10,20,40, and 80 together with six spanwise (Z) positions, three directly in line with the resonator orifices and three half-way between the orifices. Laser velocimeter measurements of the streamwise and vertical velocity were obtained and processed to yield values of the

Correlations of Rotor Wake/Airframe Interaction Measurements with Flow Visualization Data

Narayanan M. Komerath Associate Professor School of Aerospace Engineering Georgia Institute of Technology Atlanta, Georgia 30332 revised version of a paper included in the Proceedings of the 46th Annual Forum of the AHS. ABSTRACT Interaction between the aerodynamics of the rotor and the airframe causes large unsteady pressure fluctuations on rotorcraft airframes. A two-bladed rotor and a hemispherecylinder airframe model are used to study these pressure fluctuations in a wind tunnel, simulating low-speed forward flight conditions. Controlled displacement of the airframe is used to obtain finelyspaced pressure data. These are correlated with azimuth-resolved, quantitative laser sheet visualization of the dynamics of the tip vortices and the inboard vortex sheets from the rotor, along the top of the airframe. Three prominent periodic interactions are observed on the airframe surface. These are the effects of blade motion, the approach and impingement of the tip vortex, and of the vortex sheet. These interactions determine, to a large degree, the overall pressure distribution on the airframe surface, and hence the airframe unsteady airloads. The pressure signatures from the blade passage and the vortex impingement are quite different, and are explained using 2-dimensional models. NOMENCLATURE Cp Pressure coefficient based on freestream dynamic pressure CPmean (Pmean -P∞)/q∞ CPinst (Pinst -P∞)/q∞ CPsteady CP(V∞/Vsteady)2 CPunsteady Punsteady/q∞ CT Rotor thrust coefficient, T/(ρπΩ2R4) h Vertical distance between 2-D vortex core and surface H Vertical spacing between rotor hub center and airframe centerline Pmean Mean static pressure P∞ Freestream static pressure Pinst Instantaneous static pressure Punsteady Unsteady component of static pressure (Pinst Pmean) q∞ Freestream dynamic pressure R Rotor radius T Rotor Thrust U Velocity at surface beneath 2-D vortex Vblade Velocity of blade section Vsteady Far field velocity in a steady reference frame. *: This work was part of the first author's Ph.D. dissertation, performed at the School of Aerospace Engineering, Georgia Institute of Technology. This is a V∞ Tunnel free-stream speed X Streamwise distance from hub center XL Local horizontal coordinate Revised from paper at the AHS Forum, 1990. XN X co-ordinate of airframe nose

An experimental investigation of turbulent shear flow cavitation

Cavitation inception in a turbulent shear layer was studied at Reynolds numbers up to 2 × 106. Flash photography, high-speed motion pictures and holography were used to study the relation of cavitation inception to the shear-layer turbulent structure. Both spanwise and streamwise vortices were clearly visualized by the cavitation. Cavitation inception consistently occurred in the streamwise vortices and more fully developed cavitation was visible in both structures, with the streamwise cavities typically confined to the braid regions between adjacent spanwise vortices. The strength of the streamwise vortices was estimated using a Rankine vortex model, which showed that their strength was always less than 10% of that of the spanwise vortices. Measurements of fluctuating pressures were made by holographically monitoring the size of air bubbles injected into the non-cavitating shear flow. The measured pressure fluctuations had positive and negative peaks as high as 3 times the free-stream dynamic pressure, sufficient to explain cavitation inception at high values of the inception index. The occurrence of inception in the streamwise vortices of the shear layer, combined with previous reports of velocity dependence of the streamwise vortex strength, may explain the commonly observed Reynolds-number scaling of the cavitation inception index in shear flows.

Effectiveness of passive devices for axisymmetric base drag reduction at Mach 2

Experiments have been made to assess the effectiveness of passive devices for reducing base and total afterbody drag at Mach 2. The devices examined include primarily base cavities and ventilated cavities. Results show that ventilated cavities offer significant base-drag reduction and net-drag reduction of engineering value. A correlation of base- and net-drag reduction for ventilated cavities has been suggested. Typical effects of these devices on boat-tailed and flared bases have also been assessed. ^DB -pb D h L M #00 t Nomenclature = forebody (max) cross-sectional area, 490.625 mm2 = total area of ventilation = total afterbody drag coefficient, net-drag force/to^ A) = base-drag coefficient = base-pressure coefficient = forebody (max) diam, 25 mm = cavity depth = afterbody length = freestream Mach number = freestream dynamic pressure = cavity lip thickness

Planing and impacting plate forces at large trim angles

Added mass theory has been shown to give excellent agreement with experimental measurements on planing surfaces at normal planing angles [e.g. Payne, P.R. (1982, Ocean Engng 9, 515–545; 1988, Design of High-Speed Boats, Volume 1: Planning. Fishergate. Inc., Annapolis, Maryland)] and to agree exactly with more complex conformal transformations where such a comparison is possible. But at large trim angles, it predicts non-transient pressures that are greater than the free-stream dynamic pressure and so cannot be correct. In this paper, I suggest that the reason is because, unlike a body or a wing in an infinite fluid, a planing plate only has fluid on one side—the “high pressure” side. So the fluid in contact with the plate travels more slowly as the plate trim angle (and therefore static pressure) increases. This results in lower added mass forces than Munk, M. (1924) The Aerodynamic Forces on Airship Hulls (NACA TR-184) and Jones, R. T. (1946) Properties of Low-Aspect-Ratio Pointed Wings at Speeds Below and Above the Speed of Sound (NACA TR-835) originally calculated for wings and other bodies in an infinite fluid. For simplicity of presentation, I have initially considered the example of a triangular (vertex forward) planning plate. This makes the integration of elemental force very simple and so the various points are made without much trouble. But the penalty is that there seem to be no experimental data for such a configuration; at least none that I have been able to discover. But at least the equations obtained in the limits of zero and infinite aspect ratio, small trim angles (τ) and τ = 90° all agree with established concepts and the variation of normal force with trim angle looks like what we would expect from our knowledge of how delta wings behave in air. I then employed the new equation to calculate the force on a rectangular planing surface at a trim angle τ, having a constant horizontal velocity u o and a vertical impact velocity of ż. This happens to have been explored experimentally by Smiley, R. F. [(1951) An Experimental Study of Water Pressure-Distributions During Landings and Planing of a Heavily Loaded Rectangular Flat Plate Model (NACA TM 2453)] up to trim angles of τ = 45°, and so a comparison between theory and experiment is possible. The results of this comparison are encouraging, as is also a comparison with the large trim angle planing plate measurements of Shuford, C. L. [(1958) A Theoretical and Experimental Study of Planing Surfaces Including Effects of Cross Section and Plan Form (NACA Report)]. As two practical applications, I first employed the new equations to calculate the “design pressures” needed to size the plating of a transom bow on a high-speed “Wavestrider” hull. The resulting pressures were significantly different to those obtained using semi-empirical design rules in the literature. Then I used the theory to critically review data obtained from tank tests of a SES bow section during water impact to identify how the “real world” of resilient deck plating diverged from the “model world” of extreme structural rigidity.

A New Method of Modeling Underexpanded Exhaust Plumes for Wind Tunnel Aerodynamic Testing

A new method for modeling hot underexpanded exhaust plumes with cold model scale plumes in aerodynamic wind tunnel testing has been developed. The method is applicable to aeropropulsion testing where significant interaction between the exhaust and the free stream and aftbody may be present. The technique scales the model and nozzle external geometry, including the nozzle exit area, matches the model jet to free-stream dynamic pressure ratio to full-scale jet to free-stream dynamic pressure ratio, and matches the model thrust coefficient to full-scale thrust coefficient. The technique does not require scaling of the internal nozzle geometry. A generalized method of characteristic computer code was used to predict the plume shapes of a hot (γ = 1.2) half-scale nozzle of area ratio 3.2 and of a cold (γ = 1.4) model scale nozzle of area ratio 1.3, whose pressure ratio and area ratio were selected to satisfy the above criteria and other testing requirements. The plume shapes showed good agreement. Code validity was checked by comparing code results for cold air exhausting into a quiescent atmosphere to pilot surveys and shadowgraphs of model nozzle plumes taken in a static facility.

Measurement of wind effects on a 300-M concrete chimney — reference paramters

Indirect methods based on the measured pressure distribution around a chimney frequently represent the only means of determining the primary reference parameter in a chimney experiment, the free-stream dynamic pressure. But because pressures are generally measured relative to an unknown reference pressure, these methods rely on assumptions which may lead to unpredictable results. In this particular experiment assumptions were avoided by using a specially developed barometric pressure probe, suspended from a cable strung between two chimneys. Measurements are presented to show that acceptable results can be obtained by assuming the zero-crossing angle of the pressure distribution.