Showing papers on "Drag divergence Mach number published in 1994"

Prediction of ice shapes and their effect on airfoil drag

Abstract: Calculation of ice shapes and the resulting drag increases are presented for a NACA 0012 airfoil. The calculations were made using a combination of modified LEWICE and interactive boundary-layer codes for a wide range of values of parameters such as airspeed and temperature, the droplet size and liquid water content of the cloud, and the angle of attack of the airfoil. Based on experimental data, an improved correlation of equivalent sand-grain roughness was developed. Calculated ice shapes are in good agreement with experimental data for rime ice, but some differences are shown between predictions and experimental data for glaze ice. Calculated drag coefficients generally follow trends shown by the experimental data.

In-flight lift-drag characteristics for a forward-swept wing aircraft and comparisons with contemporary aircraft)

Abstract: Lift (L) and drag (D) characteristics have been obtained in flight for the X-29A airplane (a forward swept-wing demonstrator) for Mach numbers (M) from 0.4 to 1.3. Most of the data were obtained near an altitude of 30,000 ft. A representative Reynolds number for M = 0.9, and a pressure altitude of 30,000 ft, is 18.6 x 10(exp 6) based on the mean aerodynamic chord. The X-29A data (forward-swept wing) are compared with three high-performance fighter aircraft: the F-15C, F-16C, and F/A18. The lifting efficiency of the X-29A, as defined by the Oswald lifting efficiency factor, e, is about average for a cantilevered monoplane for M = 0.6 and angles of attack up to those required for maximum L/D. At M = 0.6 the level of L/D and e, as a function of load factor, for the X-29A was about the same as for the contemporary aircraft. The X-29A and its contemporaries have high transonic wave drag and equivalent parasite area compared with aircraft of the 1940's through 1960's.

The effect of compressibility on vortex pairing

Abstract: Direct numerical simulations are conducted to investigate in detail the effect of Mach number on vortex pairing in a mixing layer. The pairing process is found to be delayed at higher Mach numbers and the paths followed by the vortices change. To investigate the effect of the initial shape of the vortices a simple vortex dynamical model of pairing is constructed which accurately models pairing at low Mach numbers. Results from the model suggest that a variation in the initial shape of the vortices is not sufficient to explain the changes in the pairing process due to Mach number. Further simulations are conducted for isolated vortex pairs. There is little departure from the expected rotation rate as Mach number is increased, but strong core effects. Overall, changes in the pairing process reflect changes in the evolution of the primary instability, with vortex trajectories becoming more elongated as the Mach number is increased.

Drag Reduction for Turbulent Flow over a Projectile: Part I

Abstract: A numerical study is made to analyze the drag performance of a secant-ogive-cylinder-boattail projectile in the transonic Mach number regime between 0.91 and 1.20. To improve the projectile's performance, two drag reduction methods, boattailing and base bleed, are applied. The effectiveness of each method and the combination of both methods are studied by varying the values of parameters such as boattail angle, bleed quantity, and bleed area. The computed distributions of surface pressure coefficient of the projectile with different boattail angles are in close agreement with experimental data. Computed drag components and the total drag of the projectile are accurate by comparison with experimental data and semiempirical predictions. The optimal boattail angle for total drag reduction is predicted to be at about 5-7 deg. The method of combining boattailing and base bleed can become an effective method for total drag reduction.

Minimum-Drag Axisymmetric Bodies in the Supersonic/Hypersonic Flow Regimes

Abstract: A study of minimum-drag body shapes was conducted over a Mach range from 3 to 12. Numerical results show that power-law bodies result in low-drag shapes, where the power n = 0.69 (l/d = 3) or n = 0.70 (l/d = 5) shapes have lower drag than theoretical minimum results (« = 0.75 or 0.66, depending on the particular form of the theory). To evaluate the results, a numerical analysis was made, including viscous effects and the effect of a gas model. None of these considerations altered the conclusions. The Hayes minimum-drag body was analyzed and had a higher drag than the optimum power-law body. d i n r T rref x, y, 0 Nomenclature = drag coefficient based on the maximum cross-sectional area = skin-friction coefficient = pressure coefficient = body diameter at the base = marching plane index = unified supersonic-hypersonic similarity parameter, = body length = freestream Mach number = power-law exponent = body radius = temperature at the body surface = freestream temperature z - physical coordinates = circumferential angle

Computational Prediction of Isolated Performance of an Axisymmetric Nozzle at Mach Number 0.90

Abstract: An improved ability to predict external propulsive performance was incorporated into the three-dimensional Navier-Stokes code PAB3D. The improvements are the ability to account for skin friction and external pressure forces. Performance parameters for two axisymmetric supersonic cruise nozzle configurations were calculated to test the improved methodology. Internal and external flow-field regions were computed using a two-equation kappa-epsilon turbulent viscous-stress model. The computed thrust-minus-drag ratios were within 1 percent of the absolute level of experimental data and the trends of data were predicted accurately. The predicted trend of integrated nozzle pressure drag matched the trend of the integrated experimental pressure drag over a range of nozzle pressure ratios, but absolute drag levels were not accurately predicted.

Aerodynamic test and analysis of wrap-around fins at supersonic Mach numbers utilizing design of experiments

Abstract: Cxb =body axial force coefficient at zero angle of The aerodynamics of wrap-around fins (WAFs) are influenced by various factors, including Mach number, Icngth-to-diameter (LD) ratio, fin sweep angle, fin root chord length, and fin thickness. This paper presents the rcsults of an investigation into the effects of these factors on WAFs at supersonic Mach numbers. This effort was conducted to furthcr define WAF flight characteristics. The design of experiments (DOE) test approach was used to detcrmine the feasibility of DOE in extracting aerodynamic coefficients from spark range testing, as well as to provide a measure of cost savings. The results of the free flight data reduction and predictions arc summarized and compared to the DOE model and to engineering code predictions. These results show the effect of each factor on the tin axial force coefficient, fin normal force, and total pitching moment derivative coefficients. In addition, an analysis was completed for thc roll and side force. The results of this analysis show good agrcement with the DOE modcl and engineering prediction codes. This study also shows that DOE methodology is an efficient way to predict aerodynamic coefficients for spark range testing. -i

Preliminary Assessment of Tunnel Wall Interference in the NDA Cryogenic Wind Tunnel

Abstract: As the second stage of a preliminary cryogenic airfoil testing program, the test section interference of the National Defense Academy cryogenic tunnel was evaluated A R4 airfoil model, which has the chord length of 12 cm and the aspect ratio of 05, was tested in the range of Mach number 05 to 075, and that of Reynolds number 7 x 106 to about 11 x 10 7 The experimental results were corrected with empirical wall interference correction methods, the Barnwell-Sewall method for sidewall boundary layers, and the Blackwell method for the top and bottom walls This preliminary evaluation showed that the sidewall boundary layers dominate the tunnel wall interference of the present cryogenic tunnel, and there may be some possibility of utilizing the tunnel for performing two-dimensional airfoil tests if more precise wall interference parameters are obtained Nomenclature b = tunnel width Cp = pressure coefficient Cp = equivalent pressure coefficient c = chord length of the airfoil cn = section normal force coefficient H = shape factor M - local Mach number Mc = corrected freestream Mach number ME = averaged experimental Mach number at the wall over the airfoil MT = averaged theoretical Mach number at the wall over the airfoil Mx = tunnel freestream Mach number M^, = equivalent Mach number a = angle of attack AMB = blockage Mach number correction AMy = sequential Mach number correction AMw = sidewall Mach number correction 8* = sidewall displacement thickness Subscripts ambient condition

Hypervelocity WInd Tunnel Number 9, High Mach Number Development Program.

Abstract: : This report describes the results of the high Mach Number Development program performed at the White Oak, Maryland site of the Dahlgren Division, Naval Surface Warfare Center. The goal of this program was to expand the capabilities of the Hypervelocity Wind Tunnel Number 9 (Tunnel 9) to include operation at Mach 18. The constraints of this program involved using the existing Mach 14 setup with as little modification as necessary. There were two major areas of interest for this program, the heater and the nozzle. The required supply temperature for Mach 18 operation is above the current capabilities of the Tunnel 9 Mach 14 heater. Utilizing supercooled flow conditions lowered the required supply temperature to within the Mach 14 heater capability. The current Mach 14 nozzle was used and the throat section was replaced with a new hardware set designed to achieve the correct nozzle throat-to-exit area ratio to obtain the higher Mach number Fortyone runs were carried out in Tunnel 9 in support of this program. The Mach number capability in Tunnel 9 has been extended to Mach 16.5. For this condition the flow has a 30-inch test core with Pitot pressure deviations of-1.1% to + 1.3% and 3.5 seconds of good run time. A Mach 18 capability has also been investigated. Research efforts to achieve acceptable Mach 18 conditions are continuing.

Mach number, flow angle, and loss measurements downstream of a transonic fan-blade cascade

Abstract: : Two dimensional flow measurements of Mach number and flow angle were conducted downstream of a transonic fan-blade cascade at a mach number of 1.4 to provide baseline data for assessing the effect of vortex generating devices on the suction surface shock-boundary layer interaction. The experimental program consisted of the design and calibration of a traversing three-port pneumatic probe to measure Mach number and flow angle and initial cascade measurements to provide baseline data for the fully-mixed-out total pressure loss coefficient and flow turning angle. Similar tests are planned with the vortex generating devices installed. Comparisons with and without the vortex generating devices are needed to quantify the overall effect on the shock-boundary interaction in a transonic fan-blade passage, and to assess the potential for using vortex generating devices in military engine fans.

Development of Transonic Area-Rule Methodology

Abstract: This study presents an innovative method for area ruling in the transonic regime. The method applies a weighting function to the sonic area rule that generally accounts for the physical nature of transonic flow. In sonic flow, changes in pressure are communicated with negligible dissipation along Mach planes. As a result, drag becomes a strong function of the cross-sectional area development of the aircraft. Transonic flow has the added complexity of mixed subsonic and supersonic regions. In this flow, the communication between the aircraft fuselage and its external parts has dissipation due to embedded subsonic regions. Therefore, the sonic area rule no longer strictly applies. The new transonic area-rule methodology, described in this article, utilizes a weighting function that adjusts for the effects of the mixed flows. The shaping methods resulting from this new transonic area-ruling technique are much less severe than the standard sonic area-ruling method and require substantially less body modification. Furthermore, the new transonic area-ruling technique maintains drag rise delays that are the same as the traditional sonic area rule. Nomenclature A = stream-tube area CD = drag coefficient M = local Mach number Wm = wingtlp parameter v = velocity ACD = wave drag coefficient

An investigation of the transonic pressure drag coefficient for axi-symmetric bodies

Abstract: : This thesis investigates the pressure drag coefficient in the transonic regime over an axi-symmetric body, with a set of unique contour surfaces developed in a previous thesis. The contour surfaces were obtained by an exact solution of the small perturbation transonic equation, using the guidelines and tools developed at NPS. In this work, Computational Fluid Dynamics (CFD) was not only used to compute the afterbody contour surface, but also to investigate a conical afterbody and complete bodies, which are composed of an arbitrary forebody (ellipsoid) and a variable afterbody (contour and conical). Euler as well as Navier-Stokes flow-solvers were applied to the geometries of interest, giving Mach-number contours for viscous and inviscid flow, pressure drag coefficient magnitude, and depicting shock wave location. On the basis of these results, it can be verified that our contour surface afterbodies will decrease by 15% the peak of the pressure drag coefficient (C sub d) versus Mach number curves in the transonic regime. These results can be used to design low pressure drag surfaces for such as missiles, projectiles and aircraft engine nacelles. Transonic, Pressure, Drag, Coefficient, Axi-symmetric bodies.

Experimental Studies of Very-High Mach Number Hydrodynamics

Abstract: : We present results of experiments on very-high Mach number (> 100) shocks and very-high Mach number and Reynolds number (> 100, 106) turbulence. Such high Mach number hydrodynamics are initiated with a powerful laser pulse driver. We show that shocks created with a laser driver follow the Taylor-Sedov self-similar solution and scale via the Sachs scaling law just like shocks created by more traditional methods. In one experiment we examined laser- produced-shock solid-surface interactions and observed expected phenomena such as Marsh stems and triple points, and also measured a new phenomenon termed a blast wave decursor. In second experiment we found that shocks become unstable if they propagate through a gas which has a low adiabatic index and we measured the growth rate of the instability. In a third experiment we have shown that a high Mach number shock dramatically enhances the structure of a turbulent field through which it passes and that the shock is itself badly distorted. This result is unexpected since common wisdom has it that high Mach number shocks would self-heal as they pass through a turbulent field. Turbulence, Shock.

Mach and Load Distribution

Abstract: The choice of the Mach µkernel as an underlying environment for the LD scheme is of particular importance. A practical LD scheme should be implemented on a widely used and sufficiently sophisticated operating system, such as Mach. It is a widely used µkernel in both research and commercial communities. Its architecture supports the sophisticated enhancements that are significant for the LD design and implementation.

A numerical study of optimal drag reduction for turbulent transonic projectiles using a passive control

Abstract: The purpose of this research is to numerically study a drag reduction method—passive control of shock/boundary layer interaction, which is applied to the boattail portion of a secant-ogive-cylinder-boattail projectile in turbulent transonic flows. The flow pattern and the components of aerodynamic drag computed from numerical data are analyzed. The effectiveness of this method is studied by varying the values of parameters such as porosity distribution, maximum porosity factor and size of porous region. The conditions for optimal drag reduction are investigated and reported. The present results show that the use of this passive control method can not only reduce the boattail drag but also the base drag, and results in an additional 8% total drag reduction compared to that without the passive control technique. This passive control method can be an effective approach for the design of high-performance projectiles in the transonic regime.

Studies of turbulent. Drag reduction using riblets on a NACA 0012 aerofoil

Abstract: Results of viscous drag reduction using 3M riblets 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 stream wise 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 show13; significantly higher skin friction drag reduction with incidence compared to flat plate flows; the reduction was as high as 15% at a =6xB0;. Results of mean velocity profiles show that larger contribution to drag reduction results from the suction side of the airfoil, indicating increased effectiveness of riblets in adverse pressure gradients. Examination of turbulent intensity profiles in the wall region indicates appreciable reduction in the presence of riblets; correspondingly, the energy spectra shows reduced energy levels at low frequencies