Showing papers on "Drag divergence Mach number published in 2003"

Effect of turbulence on the drag and lift of a particle

Abstract: A direct numerical simulation (DNS) is used to study the effect of a freestream isotropic turbulent flow on the drag and lift forces on a spherical particle. The particle diameter is about 1.5–10 times the Kolmogorov scale, the particle Reynolds number is about 60–600, and the freestream turbulence intensity is about 10%–25%. The isotropic turbulent field considered here is stationary, i.e., frozen in time. It is shown that the freestream turbulence does not have a substantial and systematic effect on the time-averaged mean drag. The standard drag correlation based on the instantaneous or mean relative velocity results in a reasonably accurate prediction of the mean drag obtained from the DNS. However, the accuracy of prediction of the instantaneous drag decreases with increasing particle size. For the smaller particles, the low frequency oscillations in the DNS drag are well captured by the standard drag, but for the larger particles significant differences exist even for the low frequency components. In...

Computational Fluid Dynamics-Based Drag Prediction and Decomposition

Abstract: A method for the computation and breakdown of the aerodynamic drag into viscous and wave components is proposed. Given a numerical solution of the Reynolds averaged Navier-Stokes equations, the method, based on a Taylor's series expansion of the far-field drag expression, allows for the determination of the drag related to entropy variations in the flow. The identification of a spurious contribution, due to the numerical dissipation and discretization error of the flow solver algorithm, allows for drag computations weakly dependent on mesh size. Therefore, accurate drag evaluations are possible even on moderately sized grids. Results are presented for transonic flows around an airfoil and a wing-body configuration

Natural-Laminar-Flow Airfoil Development for a Lightweight Business Jet

Abstract: A 15% % thick, natural-laminar-e ow airfoil, the SHM-1, has been designed to satisfy requirements derived from the performance specie cations for a lightweight business jet. The airfoil was tested in a low-speed wind tunnel to evaluateitslow-speedperformance.Ae ighttestwasalsoconductedtoevaluatetheperformanceoftheairfoilathigh Reynolds numbers and high Mach numbers. In addition, a transonic wind-tunnel test was conducted to determine the drag-divergence characteristics. The design requirements, methodology, and experimental verie cation are described.

Structured and Unstructured Computations on the DLR-F4 Wing-Body Configuration

Abstract: The accuracy of the DLR structured and unstructured computational fluid dynamic (CFD) codes in predicting aircraft forces and moments on a wing-body configuration at transonic speeds is investigated. The computations form the contribution of the DLR and Airbus Deutschland to the AIAA CFD Drag Prediction Workshop in June 2001. Using a combination of a high quality grid, low levels of artificial dissipation and an advanced turbulence model, the structured code (FLOWer) was able to both qualitatively and quantitativele predict the experimentally measured drag, lift and pitching moments. In its current implementation the unstructured code (TAU) was found to be less accurate in predicting forces and moments, although qualitatively the results were good.

Wave-Drag Characteristics of an Over-the-Wing Nacelle Business-Jet Configuration

Abstract: This paper presents the wave-drag characteristics of an over-the-wing nacelle cone guration. The e ow over the wing is accelerated such that the aerodynamic interference between the nacelle and the wing is critical in the transonic e ight regime.In general, locating nacellesoverthewing causesanunfavorableaerodynamicinterference and inducesa strongshock wave,which resultsina lowerdrag-divergenceMach number. Ifthenacelleislocated at the optimum position relative to the wing, however, the shock wave can be minimized, and drag divergence occurs at a Mach number higher than that for the clean-wing cone guration. Theoretical analyses and experimental measurements demonstrate that a wave-drag reduction can be achieved by locating the nacelle front face near the shock-wave position on the wing.

Aerodynamic Drag of Engine-Cooling Airflow With External Interference

Abstract: This report examines the aerodynamic drag and external interference of engine cooling airflow Much of the report is on inlet interference, a subject that has not been discussed in automotive technical literature It is called inlet spillage drag, a term used in the aircraft industry to describe the change in inlet drag with engine airflow The analysis shows that the reduction in inlet spillage drag, from the closed front-end reference condition, is the primary reason why cooling drag measurements are lower than would be expected from free stream momentum considerations In general, the free stream momentum (or ram drag) is the upper limit and overstates the cooling drag penalty An analytical expression for cooling drag is introduced to help the understanding and interpretation of cooling drag measurements, particularly the interference at the inlet and exit Empirical thrust coefficients, which represent the interferences, are introduced as a practical representation of the interaction to the exterior pressure distribution Comparisons to experimental measurements on three vehicles are presented to illustrate the concepts

Parametric Study of Flutter for an Airfoil in Inviscid Transonic Flow

Abstract: With the use of a state of the art inviscid computational fluid dynamic harmonic balance aerodynamic-Euler-based code, a systematic, parametric investigation is presented into how the several structural parameters and freestream Mach number of a transonic flow affect the flutter characteristics of a typical two-degree-of-freedom transonic airfoil configuration. The computational efficiency of the time-linearized option of the harmonic balance aerodynamic model allows a much more thorough exploration of the parameter range than has been possible previously.

Method for reducing the drag of blunt-based vehicles by adaptively increasing forebody roughness

Abstract: A method for reducing drag upon a blunt-based vehicle by adaptively increasing forebody roughness to increase drag at the roughened area of the forebody, which results in a decrease in drag at the base of this vehicle, and in total vehicle drag.

スクラムジェットエンジンの3分力測定 1)マッハ4から8飛行条件におけるエンジン抗力

Abstract: Aerodynamic performance of scramjet engines was measured by using 0.44-m-long models in a M6.7 wind tunnel. Drags and wall pressure distributions were measured to evaluate the total pressure and friction drag in the engine internal flow. The internal drag of the models with various struts was dicussed. The internal and the external drags, the pressure and the friction drags were estimated. Consistency between the force balance measurements and the pressure measurements was examined. The internal drag obtained from the force balance agreed with that based on the wall pressure measurements.

The NASA Langley 8-Foot Transonic Pressure Tunnel Calibration

Abstract: The NASA Langley 8-Foot Transonic Pressure Tunnel is a continuous-flow, variable-pressure wind tunnel with control capability to independently vary Mach number, stagnation pressure, stagnation temperature, and humidity. The top and bottom walls of the test section are axially slotted to permit continuous variation of the test section Mach number from 0.2 to 1.2; the slot-width contour provides a gradient-free test section 50 in. long for Mach numbers equal to or greater than 1.0 and 100 in. long for Mach numbers less than 1.0. The stagnation pressure may be varied from 0.25 to 2.0 atm. The tunnel test section has been recalibrated to determine the relationship between the free-stream Mach number and the test chamber reference Mach number. The hardware was the same as that of an earlier calibration in 1972 but the pressure measurement instrumentation available for the recalibration was about an order of magnitude more precise. The principal result of the recalibration was a slightly different schedule of reentry flap settings for Mach numbers from 0.80 to 1.05 than that determined during the 1972 calibration. Detailed tunnel contraction geometry, test section geometry, and limited test section wall boundary layer data are presented.

Aerodynamic Drag and Drag Reduction

Abstract: An assessment of the role of fluid dynamic resistance and/or aerodynamic drag and the relationship to energy use in the United States is presented. Existing data indicates that up to 25% of the total energy consumed in the United States is used to overcome aerodynamic drag, 27% of the total energy used in the United States is consumed by transportation systems, and 60% of the transportation energy or 16% of the total energy consumed in the United States is used to overcome aerodynamic drag in transportation systems. Drag reduction goals of 50% are proposed and discussed which if realized would produce a 7.85% total energy savings. This energy savings correlates to a yearly cost savings in the $30Billion dollar range.

Supercritical airfoil design for future hale concepts

Abstract: This paper presents design and analysis of a new airfoil for future High-Altitude Long-Endurance (HALE) aircraft that has an operational condition at supercritical speeds. The XFOIL and MSES computational codes were used to design, modify and analyze the airfoil. The airfoil has enough thickness and performance to meet the requirements set for one of the AFRL SensorCraft concepts; a joined-wing configuration with diamondshape in planform and front views. This SensorCraft concept’s geometry and operational altitudes and speeds were used to determine the airfoil design conditions. Sensitivity studies were carried out to investigate the effects of Reynolds number and Mach number, along with boundary layer transition parameters. The airfoil has a drag bucket over a large range of lift coefficient. Boundary layer transition location is at about 60% chord upper and 70% chord lower surface, and characterized by a laminar separation bubble, which decreases in size with increases in angle of attack. Further work needs to be performed to validate the design with experiments. NOMENCLATURE AR = Aspect ratio a = freestream speed of sound c = chord CL = airfoil lift coefficient CD = airfoil drag coefficient CM = airfoil moment coefficient CP = surface pressure coefficient L = lift force D = drag force M = freestream Mach number, V/a MD = drag divergence Mach number M = reduced Mach number, L

Influence of Loading Distribution on the Performance of Transonic HP Turbine Blades

Abstract: Midspan measurements were made in a transonic wind tunnel for three HP turbine blade cascades at design incidence. The baseline profile is the midspan section of a HP turbine blade of fairly recent design. It is considered mid-loaded. To gain a better understanding of blade loading limits and the influence of loading distributions, the profile of the baseline airfoil was modified to create two new airfoils having aft-loaded and front-loaded pressure distributions. Tests were performed for exit Mach numbers between 0.6 and 1.2. In addition, measurements were made for an extended range of Reynolds numbers for constant Mach numbers of 0.6, 0.85, 0.95 and 1.05. At the design exit Mach number of 1.05, the aft-loaded airfoil showed a reduction of almost 20% in the total pressure losses compared with the baseline airfoil. However, it was also found that for Mach numbers higher than the design value the performance of the aft-loaded blade deteriorated rapidly. The front-loaded airfoil showed generally inferior performance compared with the baseline airfoil.Copyright © 2003 by ASME

A Relationship Between Wave Drag and Induced Drag

Abstract: A formulation for the induced drag of a wing in subsonic or transonic flow is derived from entropy considerations. This approach shows how wave drag and induced drag are related. The new formulation is cast in a form similar to that used in the classic induced drag derivation thus allowing a theoretical comparison of the two approaches. If there are no shock waves in the flow the two formulations agree theoretically only in the case of an elliptic wing loading, although calculations indicate that the quantitative difference may be relatively small. If shock waves are present they can increase or decrease the induced drag leading to the idea of a reduction in the sum of induced and wave drag by a judicious tailoring of the flow over the wing.

Supersonic and Hypersonic Minimum Drag for Bodies of Revolution

Abstract: Forebody drag measurements have been acquired on a variety of nose geometries with fineness ratios ranging from 2 to 4, over an angle of attack range from 0 to 10 degrees, and at Mach numbers of 3, 4.5, and 6. Nose geometries tested included a tangent ogive, von Karman, power-series and cone. Test data indicates that the 0.7 power-series nose out-performs other geometries over the majority of the test conditions. Test data is compared to predictions made using the USAF Missile DATCOM and the Aeroprediction Inc./NSWC AP02 codes. In general both codes predict the drag within about 5-percent, however it is shown that the AP02 code more accurately estimates angle of attack effects on axial force. Nomenclature CAF Forebody axial force coefficient (body-fixed axis system) CAU Uncorrected forebody axial force coefficient (body-fixed axis system) CD Drag coefficient (wind axis system) Lref Reference length, diameter of model (2.214 in.=1 caliber) M Freestream Mach number q Freestream dynamic pressure, psi Sref Reference area based on reference length, (3.850 in) α Missile angle of attack, deg *Aerospace Engineer, Senior Member AIAA ** Aerospace Engineering Intern This paper is declared a work of the U.S. Government and is not subject to copyright protection in the United States. Introduction The development of a tube-launched, hypervelocity missile that operates at sea level presents unique challenges in missile and projectile configuration design. From an aerodynamic standpoint, the minimization of total integrated drag during flight is critical to obtain maximum performance in range and lethality. As part of an extensive missile configuration trade study, public domain literature spanning over 50 years have been reviewed to determine if there is a nose planform that has minimum drag over the intended operational flight envelope (Mach 0 to 6). While the literature search provided a wealth of data, there was an apparent lack of directly comparable data. As a result, a study was undertaken to obtain forebody drag data on a wide range of nose planform shapes in an effort to identify the lowest drag nose planform. The objectives of this paper are to present recently acquired test data for a variety of nose planform geometries and to present comparisons of that data with predictions from engineering design codes. Historical Review In 1951, Perkins and Jorgensen published data indicating that the theoretical shapes for minimum diameter or given diameter and volume derived by von Karman and Haack do not have less drag than all other possible shapes having identical values of the same parameters. Their data covered a Mach range of 1.24 to 3.67 and also indicated that for a fixed fineness ratio of 3, the fore drag is reduced somewhat by small degrees of blunting, however for fixed cone angle, blunting always increased drag. In 1957, Eggers, Resnikoff and Dennis published data for various nose shapes over a Mach range of 2.73 to 6.28 that indicated that the 3⁄4 power-law body has the minimum forebody drag as re-presented in Figure 1. Figure 2 (also extracted from reference 3) indicates that modified Newtonian Impact theory predicts that the 0.7 power-law body may have the minimum drag for 3-caliber nose lengths. Perkins, Jorgensen, and Sommer, published data in 1958 that indicated that the paraboloid of revolution (1/2 power21st Applied Aerodynamics Conference 23-26 June 2003, Orlando, Florida AIAA 2003-3417 This material is declared a work of the U.S. Government and is not subject to copyright protection in the United States. 2 American Institute of Aeronautics and Astronautics law nose) had the least foredrag below M=1.5 and the hypersonic optimum shape (3/4 power-law nose) had the least foredrag above M=1.5. This report also presented data for tangent ogive, cone, power-series, ellipsoid, and LD, L-V, & D-V Haack noses planform geometries. Peckham conducted experiments on power-law bodies in 1965 and noted that minimum pressure drag is obtained by accepting higher pressures on a relatively small area of large slope near the nose, with a consequent reduction in pressure on a relatively large area towards the base of the nose. Building upon the work of Eggers, Resnikoff and Dennis, Miele uses Newtonian impact theory to show that the minimum drag for given constant diameter and length is achieved by a 0.75 power-law nose, and the minimum drag for given constant diameter and wetted area is achieved by a cone or 1.00 power-law nose. In 1966, Spencer and Fox published data at Mach 10 for a family of power-law bodies having values of the exponent of 0.25, 0.50, 0.66, 0.75, and 1.00, and a theoretical hypersonic minimum-wave-drag body. This study acquired data on 12 different body shapes with circular or elliptical cross-sections. The effects of increasing ellipticity for a given power-law body or the minimum-wave-drag body resulted in an almost constant incremental increase in maximum L/D, independent of body longitudinal contour. In 1989, Wesby and Regan presented hypersonic drag data for a series of 2-caliber power-law bodies at Mach numbers of 9.84 and 12.8, in the transitional rarefied flow regime (Reynolds numbers ranging from 1936 to 350000) that indicates that 0.75 is the exponent for minimum drag, and that ~0.64 is the exponent for minimum drag per unit volume given a fixed fineness ratio. In 1990, Mason and Lee published results from a numerical study that indicated that the minimum drag power-law body would have a 0.70 exponent. Experimental Test In the current study, a series of nose geometries were tested at Mach numbers of 3.0, 4.5, and 6.0 to characterize the drag. Tests were conducted at the Lockheed Martin Missiles and Fire Control (LMMFC) High Speed Wind Tunnel (HSWT) and at the NASA Langley Research Center (LaRC) Mach 6 facilities. Table 1 summarizes the test conditions. The model was 2.214 inches in diameter and was instrumented with a 6-component internal strain-gage balance housed inside a 4.719-caliber cylindrical afterbody. Figure 3 and Figure 4 depict the model hardware and nose geometry. Four pressure ports were used to measure base pressure. Table 2 presents the axial force balance accuracies in coefficient form for the given test conditions. Experimental Results Figure 5 presents a summary of the Forebody Drag Coefficient at zero Angle-of-Attack for various nose geometries on a standard cylindrical afterbody. This figure clearly illustrates the benefits of a high fineness ratio nose. It also clearly indicates that within a given fineness ratio, the power-series (specifically the 0.7 power-series) nose has the lowest planform wave drag for the majority of the test conditions. The Missile DATCOM predicted relative magnitudes of the drag components for a 3-caliber, 0.70 power-series nose are presented in Figure 6. This data is also for a clean, nose+body configuration. The significance of base drag in the subsonic through supersonic region is clearly shown for this optimized nose planform case. Base drag may be reduced by using a sustainer motor. If, however, significant base area is present, then the nozzle flow can act as an ejector to further lower the base pressure thereby increasing the base drag. However, since this base pressure acts only over the base minus nozzle area, the over all base pressure will be reduced in the majority of the cases. Comparisons with Design Code Predictions Figure 7 and 8 compare AP02 and Missile DATCOM Axial Force predictions, respectively, with wind tunnel test data at Mach 3 conditions. Here it is seen that both codes predict the zero lift drag within about 5%, however AP02 tends to more accurately predict the angle of attack effects. Figure 9 compares the measured Uncorrected Axial Force and the measured Forebody axial force to the values predicted by AP02 and Missile DATCOM for a 2-caliber, 0.70 power-law nose planform. This figure indicates that AP02’s modeling of angle of attack effects on base pressure is the likely reason that AP02 tends to better capture the AoA effects. Figure 10 presents the predicted axial force components (wave, skin friction, and base pressure) for a 2-caliber, 0.70 power-law nose planform. This figure indicates that the most significant difference between the predicted components below 10-degrees AoA, is the AoA dependency of base pressure as modeled in AP02. Figure 11, 12, and 13 present additional comparisons of AP02 and Missile DATCOM with test results. Table 3 through Table 6 present a summary of the codes predictive accuracies for both CAU and CAF. Here it is seen that AP02 tends to be slightly more accurate than Missile DATCOM, but both tend to predict body-alone drag within about 10 percent.

Progress in plasma assisted drag reduction technology

Abstract: Using characteristics of plasma to reduce aerodynamic drag of vehicles is a new concept of reducing drag. Some recent research results in wind tunnel test and CFD indicated an obvious effect of plasma assisted drag reduction. This paper presented the development of plasma assisted drag reduction technology abroad including reducing aerodynamic drag by a counter plasma jet, plasma-based boundary layer flow control and drag reduction, reducing drag by local applying energy point sources. The fundamental principle of the plasma drag reduction and its key technologies are discussed. Finally, some research direction of plasma drag reduction technology in China are suggested.

Using Drag to Hover

Abstract: Unlike a helicopter, an insect can, in theory, use both lift and drag to stay aloft. Here we show that a dragonfly uses mostly drag to hover by employing asymmetric up and down strokes. Computations of a family of strokes further show that using drag can be as efficient as using lift at the low Reynolds number regime appropriate for insects.

Comparison of Computational and Experimental Results for a Supercritical Airfoil

Abstract: A computational investigation was performed to study the flow over a supercritical airfoil model. Solutions were obtained for steady-state transonic flow conditions using a thin-layer Navier-Stokes flow solver. The results from this computational study were compared with time-averaged experimental data obtained over a wide Reynolds number range at transonic speeds in the Langley 0.3-Meter Transonic Cryogenic Tunnel. Comparisons were made at a nominal Mach number of 0.72 and at Reynolds numbers ranging from 6 x10 super6 to 35 x 10 super6.

Wall-cavity acoustic Green’s function at low mach number

Abstract: Analytical approximations are developed for the low Mach number, aeroacoustic Green's function for a rectangular or circular cylindrical open cavity in a plane, rigid wall. The formulae can be used to predict the sound radiated into the main flow from a knowledge of the hydrodynamic flow near the cavity. At low Mach numbers the sound is a small by-product of the main flow, whose hydrodynamic properties can first be determined from observation or from a numerical treatment of the incompressible Navier-Stokes equations. Detailed predictions are made of the lowest order, open cavity resonance frequencies, and it is shown how a resonance is excited by the unsteady drag, and also by the lift or drag force experienced by a small bluff body placed in the flow close to the cavity. The cavity resonance frequencies are complex, with imaginary parts depending primarily on radiation damping, which can be sufficiently large for a shallow, open cavity, that a distinct resonance peak is absent from the acoustic spectrum...

Improved Numerical Calculation of the Airfoil Transonic Drag Applied Within a Zonal Flowfield Modeling Concept

Abstract: Very high cost efficiency of the flight is a crucial requirement specially in the contemporary commercial airplane design. Beside the low engine fuel consumption, advanced aerodynamics is another dominant factor which must be satisfied to fulfill this request. Many of these aircraft cruise at speeds slightly lower than the speed of sound, so their lifting surfaces and corresponding airfoils must be optimized primarily for this domain. One of the first steps in that process is selection or even design of the customized airfoils for the particular wing and other lifting surfaces that will produce the least possible shock wave drag in cruising flight. Nowadays the numerical airfoil optimization is very important part in that process. Algorithm presented in this paper enables the numerical calculation of wave drag both for the existing and the airfoils designed specially for a certain aircraft, and it is primarily aimed for use in the operational aircraft design. This algorithm is fairly simple and very reliable, which has been proven by comparing it’s results, obtained through the computer program Tranpro, with the experimental results for airfoils tested at several most competent aeronautical institutions throughout the world.

Aerodynamic Characteristics and Pressure Distributions for an Executive-Jet Baseline Airfoil Section

Abstract: A wind tunnel test of an executive-jet baseline airfoil model was conducted in the adaptive-wall test section of the NASA Langley 0.3-Meter Transonic Cryogenic Tunnel. The primary goal of the test was to measure airfoil aerodynamic characteristics over a wide range of flow conditions that encompass two design points. The two design Mach numbers were 0.654 and 0.735 with corresponding Reynolds numbers of 4.5 x 10 super 6 and 8.9 x 10 super 6 based on chord, respectively, and normal-force coefficients of 0.98 and 0.51, respectively. The tests were conducted over a Mach number range from 0.250 to 0.780 and a chord Reynolds number range from 3 x 10 super 6 to 18 x 10 super 6. The angle of attack was varied from -2 degrees to a maximum below 10 degrees with one exception in which the maximum was 14 degrees for a Mach number of 0.250 at a chord Reynolds number of 4.5 x 10 super 6. Boundary-layer transition was fixed at 5 percent of chord on both the upper and lower surfaces of the model for most of the test. The adaptive-wall test section had flexible top and bottom walls and rigid sidewalls. Wall interference was minimized by the movement of the adaptive walls, and the airfoil aerodynamic characteristics were corrected for any residual top and bottom wall interference.

Experiments in a Transonic Shock Tube at Subsonic Mach and High Reynolds Numbers

Abstract: Modern civil transport aircraft cruise in the high transonic velocity region near the speed of sound. Mach numbers of up to 0.86 and Reynolds numbers based on the chord length of the wing of up to Rec = 50 × 106 and higher are operational for jet planes like the Airbus A340–600 and the Boeing 747–400. Even bigger aircraft are in the conceptual or design phase with higher passenger capacities e.g. the A380, resulting in larger wing chords to lift the increasing aircraft weights in order to accomodate more passengers or freight per aircraft. Another market trend tends to higher cruising Mach numbers to shorten flight time. So flight Reynolds and Mach numbers will increase in the future.

엔진화염에 따른 천음속 유도탄의 항력 평가

Abstract: Accurate assessment of the effect of jet plume on the boattail pressure drag of transonic airbreathing missiles is very important to reduce drag and to satisfy the flight range and the required maneuver. Numerical results of drag analysis for boattail and base pressures due to jet plume are presented considering the turbulence modeling. Drag assessment due to the size of jet plume, the conditions of the exhaust gas, the configurations of the boattail, and transonic mach numbers is included.