Showing papers on "Drag divergence Mach number published in 1997"

Drag Prediction and Transition in Hypersonic Flow

Abstract: This paper discusses progress on issues such as instability studies, nose-bluntness and angle-of-attack effects, and leading-edge-contamination problems from theoretical, computational, and experimental points of view. Also included is a review of wind-tunnel and flight data, including high-Re flight transition data, the levels of noise in flight and in wind tunnels, and how noise levels can affect parametric trends. A review of work done on drag accounting and the role of viscous drag for hypersonic vehicles is also provided.

A parametric study of Mach reflection in steady flows

Abstract: The flow fields associated with Mach reflection wave configurations in steady flows are analysed, and an analytical model for predicting the wave configurations is proposed. It is found that provided the flow field is free of far-field downstream influences, the Mach stem heights are solely determined by the set-up geometry for given incoming-flow Mach numbers. It is shown that the point at which the Mach stem height equals zero is exactly at the von Neumann transition. For some parameters, the flow becomes choked before the Mach stem height approaches zero. It is suggested that the existence of a Mach reflection not only depends on the strength and the orientation of the incident shock wave, as prevails in von Neumann's three-shock theory, but also on the set-up geometry to which the Mach reflection wave configuration is attached. The parameter domain, beyond which the flow gets choked and hence a Mach reflection cannot be established, is calculated. Predictions based on the present model are found to agree well both with experimental and numerical results.

Comparative Results From a CFD Challenge Over a 2D Three-Element High-Lift Airfoil

Abstract: A high-lift workshop was held in May of 1993 at NASA Langley Research Center. A major part of the workshop centered on a blind test of various computational fluid dynamics (CFD) methods in which the flow about a two-dimensional (2D) three-element airfoil was computed without prior knowledge of the experimental data. The results of this ''blind'' test revealed: 1. The Reynolds Averaged Navier-Stokes (RANS) methods generally showed less variability among codes than did potential/Euler solvers coupled with boundary-layer solution techniques. However, some of the coupled methods still provided excellent predictions. 2. Drag prediction using coupled methods agreed more closely with experiment than the RANS methods. Lift was more accurately predicted than drag for both methods. 3. The CFD methods did well in predicting lift and drag changes due to changes in Reynolds number, however, they did not perform as well when predicting lift and drag increments due to changing flap gap. 4. Pressures and skin friction compared favorably with experiment for most of the codes. 5. There was a large variability in most of the velocity profile predictions. Computational results predict a stronger slat wake than measured suggesting a missing component in turbulence modeling, perhaps curvature effects.

Design and experimental results for the S814 airfoil

Abstract: A 24-percent-thick airfoil, the S814, for the root region of a horizontal-axis wind-turbine blade has been designed and analyzed theoretically and verified experimentally in the low-turbulence wind tunnel of the Delft University of Technology Low Speed Laboratory, The Netherlands. The two primary objectives of high maximum lift, insensitive to roughness, and low profile drag have been achieved. The constraints on the pitching moment and the airfoil thickness have been satisfied. Comparisons of the theoretical and experimental results show good agreement with the exception of maximum lift which is overpredicted. Comparisons with other airfoils illustrate the higher maximum lift and the lower profile drag of the S814 airfoil, thus confirming the achievement of the objectives.

Viscous airfoil computations using local preconditioning

Abstract: A local preconditioning technique for compressible flows at low Mach number is examined in the context of turbulent flows over airfoils. The preconditioning technique is implemented in an approximately-factored algorithm for the steady thin-layer Navier-Stokes equations. Parametric studies are presented showing trade-offs between accuracy and convergence rates. Excellent convergence rates and solution accuracy are demonstrated for freestream Mach numbers as low as 0.01. Improvements in accuracy, especially in drag prediction, are seen for Mach numbers as high as 0.4. Thus the present approach is useful in the computation of low-speed airfoil flows, including those typically encountered during take-off and landing of aircraft.

Drag coefficient and velocity of rise of a single collasping two-phase bubble

Abstract: The velocity of rise and the drag of a single vapor bubble collapsing in another immiscible liquid were measured experimentally. During the process of collapse, the dispersed-phase vapor bubble was transformed to a two-phase bubble with condensate accumulating at the rear of the two-phase bubble and vapor at its top. Such a configuration of a two-phase bubble is commonly known as a drobble. Experimental data for the six pairs of liquids covered a range of drobble (two-phase bubble) Reynolds numbers from 0.003 to 3,000. Two regimes of drobble movement were encountered. In the first regime (Re < 100), the drobble maintained its sphericity, and the observed drag was less than the solid-sphere drag predicted by the established solid sphere or Hadamard et al. fluid-sphere drag models. In the second regime, the drobble was deformed and oscillated; the observed drag departed suddenly from predictions of spherical models and increased with increasing Reynolds numbers. The critical Reynolds number covered a range from 100 to 1,000. Empirical models for drobble velocity and drag coefficient are derived.

Optimization of an Advanced Business Jet

Abstract: A new business jet is optimized to fly significantly faster than most current production aircraft while operating from relatively short runways. This new airplane is required to accommodate eight passengers in a double-club arrangement and to carry six passengers for 2800 n mile at a Mach number between 0.81-0.85. Two aircraft optimization codes are used here to ensure the validity of the design results and to identify errors in the analysis methods. These codes include the aircraft analysis methods necessary to evaluate the aircraft performance over an entire mission and optimization routines that enable the development of a family of optimum configurations. The design objective, empty weight, is shown to change approximately 1% between 30-40 deg of wing sweep at a Mach number of 0.81. At a fixed wing sweep of 31.5 deg and a Mach number of 0.81, the empty weight decreases less than 3.5% when the wing's thickness-to-chord ratio is increased from 0.10 to 0.14. A study of the design's sensitivity to Mach number indicated that the optimum empty weight and wing thickness began to change rapidly between the Mach numbers of 0.83-0.85.

Determination of Drag From Three-Dimensional Viscous and Inviscid Flowfield Computations

Abstract: A momentum balance approach is used to extract the drag from flowfield computations for wings and wing/bodies in subsonic/transonic flight. The drag is decomposed into vorticity, entropy, and enthalpy components which can be related to the established engineering concepts of induced drag, wave and profile drag, and engine power and efficiency. This decomposition of the drag is useful in formulating techniques for accurately evaluating drag using computational fluid dynamics calculations or experimental data. A formulation for reducing the size of the region of the crossflow plane required for calculating the drag is developed using cut-off parameters for viscosity and entropy. This improves the accuracy of the calculations and decreases the computation time required to obtain the drag results. The improved method is applied to a variety of wings, including the M6, W4, and Ml65 wings, Lockheed Wing A, a NACA 0016 wing, and an Elliptic wing. The accuracy of the resulting drag calculations is related to various computational aspects, including grid type (structured or unstructured) , grid density, flow regime (subsonic or transonic), boundary conditions, and the level of the governing equations (Euler or Navier-Stokes ). The results show that drag prediction to within engineering accuracy is possible using computational fluid dynamics, and that numerical drag optimization of complex aircraft configurations is possible.

Leading edge channel for enhancement of lift/drag ratio and reduction of sonic boom

Abstract: The leading edges of wings, nose assemblies, tails, fins, struts, and other components of aircraft, atmospheric entry vehicles and missiles in which the leading edge is blunted and the flight Mach number is supersonic, are provided with passive airflow channel, resulting in significantly reduced wave drag and total drag, significantly increased lift-to-drag ratio, and reduced sonic boom.

Zero-lift drag characteristics of afterbodies with a square base

Abstract: Results of zero-lift drag characteristics of afterbodies with a square base relevant to missile and projectile applications at several transonic Mach numbers and Mach 2 are presented. These afterbodies, at zero incidence, generate vortical flows upstream of the base like that on the lee side of a delta wing at incidence. The measurements made consisted primarily of afterbody drag using a balance and base pressure and extensive surface flow visualization studies were carried out to infer features associated with vortex flows. Results of base pressure, boat-tail profile drag, and total afterbody drag are presented and compared with results from the axisymmetric counterpart involving circular arc and conical boat tailing having the same base area. Some aspects of the flow features on square-base afterbodies are discussed as well.

Effect of riblets on axisymmetric base pressure

Abstract: Experiments have been performed assessing riblet effects on axisymemtric base pressure at low speeds. The tests were conducted in the 0.91-m-diam. low-speed wind tunnel at a freestream velocity of 20 m/s. It is shown that, with a large-scale or massive separation, riblets do not decrease the base drag at low speeds; on the other hand, there is a progressive increase in base drag with h+ and the drag penalty is about 8 percent for the optimized drag reducing riblet. Riblets may provide a small pressure or base drag reduction on streamlined bodies with a sharp trailing edge like an airfoil or turbine blade, because the effects of viscous-induced displacement thickness will be lower on the riblet surface.

Blade profile for helicopter blade

Abstract: PROBLEM TO BE SOLVED: To increase the drag divergence Mach number and maximum lift coefficient and reducer the noise level by defining a blade profile at the thickness ration of a specific range while the blade profile with the specific thickness ratio defined with the upper and lower faces of the blade profile by the specific coordinate system and defined with the front edge shape of the blade profile by the specific front edge radius and circle enter is used as a reference. SOLUTION: The blade profile with the thickness ratio of 8% practically defined with the upper and lower faces of the blade profile by the coordinate system expressed by the table and practically defined with the front edge shape of the blade profile by the front edge radius of r/C=0.00844 and the circle center of X/C=0.00842, Y/C=0.00064 is used as a reference, and a blade profile is defined at the thickness ratio in the range of 5-11%, where X is the distance from front edge to the rear edge of the blade profile along the chord line, C is length in the chord direction of the cross section of the blade profile, Yup is the distance from the chord line to the upper face, Yelow is the distance from the chord line to the lower face, and (r) is the front edge radium. The drag divergence Mach number and maximum lift coefficient can be increased, and the noise level can be reduced.

Simulation of Three-Dimensional Turbulent Flows around an Ahmed Body-Evaluation of Turbulence Models-

Abstract: A numerical simulation has been carried out for three-dimensional turbulent flows around an Ahmed body. The Reynolds-averaged Navier-Stokes equation is solved with the SIMPLE method in general curvilinear coordinates system. Several k-.epsilon. turbulence models with two convective difference schemes are evaluated for the performance such as drag coefficient, velocity and pressure fields. The drag coefficient, the velocity and pressure fields are found to be changed considerably with the adopted k-.epsilon. turbulence models as well as the finite difference schemes. The results of simulation prove that the RNG k-.epsilon. model with the QUICK scheme predicts fairly well the tendency of velocity and pressure fields and gives more reliable drag coefficient. It is also demonstrated that the large difference between simulations and experiment in the drag coefficient is due to relatively high predicted values of pressure drag from vertical rear end base.

Base drag reduction due to riblets on a gaw (2) airfoil

Abstract: Results from an experimental study to determine the effect of riblets on the base drag of a GAW(2) airfoil at low speeds are presented. Riblet films (3M Co.,USA) of groove height 0.076 mm and 0.152 mm were employed in the study. Measurements made consisted of surface pressure distributions, base pressure and total drag using the wake survey method in the incidence range of -2 to 6 deg. Results show that riblets provide base drag reduction of about 0.7 % of the total drag, in addition to significant viscous drag reduction.

Supersonic Transitional Airfoil Shapes of Minimum Drag

Abstract: Feasibility of a procedure incorporating transition considerations in optimizing total drag has been demonstrated. The Schittkowski algorithm was modie ed to demonstrate the approach on two-dimensional airfoils. Cubic spline basis functions were used to describe the airfoils, and total drag (wave 1 friction) was minimized under the constraint of a e xed airfoil area. Reynolds numbers were assumed such that the airfoil was transitional, generally with the forward portion laminar and the aft turbulent. The transition locus was calculated using the fast transition prediction module, which provides rapid computation of the Tollmien ‐ Schlichting wave amplie cation factor N and estimates transition point by the e N method. The laminar friction drag was evaluated using self-similar solutions of the boundary-layer equations. The friction drag for the turbulent portion was computed assuming a one-seventh velocity proe le. With this framework, the total drag was a function of the functional xtr, the streamwise location of transition. As a validation of the method, the algorithm gave the correct global optimum for the inviscid case, which is the parabolic arc proe le. For the viscous case, signie cantly different locations of the maximum thickness led to only small differences in the minimum total drag. In addition, the drag reductions from the optimal inviscid parabolic proe le were about 10%. Convergence with respect to the number of spline knots was achieved. Important challenges were met to reduce the effect of truncation errors in the numerical approximation of differentiations such as those used in the evaluation of the Hessian matrix. Despite these dife culties, generalization of the approach to treat ine nite yawed and swept wings accounting for suction, crosse ow instabilities, and vortex drag appears feasible.

Shape of a helicopter blade

Abstract: The upper surface and lower surface of a rotor blade of a helicopter are defined by the above coordinate system. The shape of the leading edge of the blade is defined by the above leading edge radius and circle center. With these shape definitions, the drag divergence Mach number Mdd and the maximum lift coefficient Clmax can be increased and the noise leved reduced.

Supersonic Transitional Airfoil Shapes of Minimum Drag

Abstract: Feasibility of a procedure incorporating transition considerations in optimizing total drag has been demonstrated. The Schittkowski algorithm was modified to demonstrate the approach on two-dimension al airfoils. Cubic spline basis functions were used to describe the airfoils, and total drag(=wave+friction) was minimized under the constraint of fixed airfoil area. Reynolds numbers were assumed such that the airfoil was transitional, generally with the forward portion laminar and the aft turbulent. The transition locus was calculated using the transition prediction module (TPM), which provides fast computation of the TollmienSchlichting wave amplification factor N and estimates transition point by the eN -method. The laminar friction drag was evaluated using self-similar solutions of the boundary layer equations. The friction drag for the turbulent portion was computed assuming a 1/7 velocity profile. With this framework, the total drag was a function of the functional, xtr, the streamwise location of transition. As a validation of the method, the algorithm gave the correct global optimum for the inviscid case, which is the parabolic arc profile. For the viscous case, significantly different locations of the maximum thickness location led to only small differences in the minimum total drag. In addition, the drag reductions from the optimal inviscid parabolic profile were about 10%. Convergence with respect to the number of spline knots was achieved. Important challenges were met to reduce the effect of truncation errors in the numerical approximation of differentiations such as those used hi the evaluation of the Hessian matrix. Special flexibility in the cutoff criteria for the iterations had to be built into the optimization routines since demanding standards for convergence could not be met. Other issues regarding noisy initial iterates (from errors in geometry, for example) propagating through the iterations as well as round-off error will need to be addressed in future effort. Despite these difficulties, generalization of the approach to treat infinite yawed and swept wings accounting for suction, cross-flow instabilities and vortex drag appears feasible.

Flowfield study over a bulbous payload shroud in transonic and low supersonic mach number

Abstract: Numerical simulation has been carried out to investigate turbulent compressible flow over a bulbous payload shroud at zero incidence in the Mach number range of 0.8 - 3.0 and Reynolds number range of 3.314x107/m - 4.682x107 /m. The numerical code solves the Reynolds averaged Navier-Stokes equations using finite-volume technique in conjunction with multistage Runge-Kutta time-stepping scheme. Closure of these equations is achieved using Baldwin-Lomax turbulence model. Comparisons have been made with experimental results such as schlieren pictures, position of terminal shock on forebody cylinder and surface pressure distribution. They are found in good agreement. A separated flow on the cylinder is observed between Mach number 0.8 - 0.9. The location of the terminal shock is found as a nonlinear function of freestream Mach number. The separation region on the boattail is estimated at various value of Mach numbers.