Showing papers on "Drag divergence Mach number published in 1991"

Drag balance for hypervelocity impulse facilities

Abstract: A new technique is described for measuring drag with 100-/*s rise time on a nonlifting model in a free piston shock tunnel. The technique involves interpretation of the stress waves propagating within the model and its support. A finite element representation and spectral methods are used to obtain a mean square optimal estimate of the time history of the aerodynamic loading. Thus, drag is measured instantaneously and the previous restriction caused by the mechanical time constant of balances is overcome. The effectiveness of the balance is demonstrated by measuring the drag on a cone with 15-deg semivertex angle in nominally Mach 5.6 flow with stagnation enthalpies from 2.6 to 33 MJ/kg. Measurement repeatability of about 10% is achieved.

Airfoil design method using the Navier-Stokes equations

Abstract: An airfoil design procedure is described that was incorporated into an existing 2-D Navier-Stokes airfoil analysis method. The resulting design method, an iterative procedure based on a residual-correction algorithm, permits the automated design of airfoil sections with prescribed surface pressure distributions. The inverse design method and the technique used to specify target pressure distributions are described. It presents several example problems to demonstrate application of the design procedure. It shows that this inverse design method develops useful airfoil configurations with a reasonable expenditure of computer resources.

Diabatic supersonic flows of dense gases

Abstract: By adopting the van der Waals gas model the behavior of n‐octane with addition or removal of heat in inviscid, steady, one‐dimensional, supersonic flow is investigated. First, the given heat distribution is added in a constant area flow. Second, the effects in variable cross‐sectional flows (nozzles) are shown for the additional condition of constant values of the Mach number or the static pressure, and for a given heat addition. As known from the isentropic flow of real gases, several gas‐dynamic effects are inverted depending on the fundamental derivative Γ. Heat addition in a constant area duct at conditions near the thermodynamical critical point may cause an increase of the Mach number in transonic/supersonic flow, or a decrease in subsonic flow, which is the opposite behavior of a perfect gas. For heat addition at M=const a singularity is possible for Γ<0 in transonic/supersonic flow and for Γ≲1 in subsonic flow. Considering constant Mach number flows, the static pressure goes up for high values of Γ. Isobaric heat addition increases the Mach number for densities below the thermodynamical critical point. Perfect gas flow exhibits the opposite reaction in both cases without any singularity. Compared with the flow of a perfect gas (γ=1.4), in the dense gas the influence of the same rate of heat addition becomes much stronger.

Mach 2 and Mach 3 mixing and combustion in scramjets

Abstract: Experimental and theoretical studies are being conducted to explore techniques to enhance mixing in scramjet combustors using parallel fuel injection from the base of swept and unswept wall-mounted ramps. Parallel injection may be useful in high speed scramjets due to the thrust contributed by the momentum of expanding fuel that has been heated in the vehicle cooling cycle. The experiments reported herein were conducted using Mach 2 and 3 combustor inlet conditions. Supporting computational and cold flow studies indicated that the observed enhanced mixing for the swept ramp configuration is primarily due to the substantially higher degree of vorticity and entrainment generated by the swept trailing edges.

Aerodynamic characteristics of a rotorcraft airfoil designed for the tip region of a main rotor blade

Abstract: A wind tunnel investigation was conducted to determine the 2-D aerodynamic characteristics of a new rotorcraft airfoil designed for application to the tip region (stations outboard of 85 pct. radius) of a helicopter main rotor blade. The new airfoil, the RC(6)-08, and a baseline airfoil, the RC(3)-08, were investigated in the Langley 6- by 28-inch transonic tunnel at Mach numbers from 0.37 to 0.90. The Reynolds number varied from 5.2 x 10(exp 6) at the lowest Mach number to 9.6 x 10(exp 6) at the highest Mach number. Some comparisons were made of the experimental data for the new airfoil and the predictions of a transonic, viscous analysis code. The results of the investigation indicate that the RC(6)-08 airfoil met the design goals of attaining higher maximum lift coefficients than the baseline airfoil while maintaining drag divergence characteristics at low lift and pitching moment characteristics nearly the same as those of the baseline airfoil. The maximum lift coefficients of the RC(6)-08 varied from 1.07 at M=0.37 to 0.94 at M=0.52 while those of the RC(3)-08 varied from 0.91 to 0.85 over the same Mach number range. At lift coefficients of -0.1 and 0, the drag divergence Mach number of both the RC(6)-08 and the RC(3)-08 was 0.86. The pitching moment coefficients of the RC(6)-08 were less negative than those of the RC(3)-08 for Mach numbers and lift coefficients typical of those that would occur on a main rotor blade tip at high forward speeds on the advancing side of the rotor disk.

Tip vortex/airfoil interaction for a low reynolds number canard/wing configuration

Abstract: The effects of the vortical wake shed by a finite span canard on a low Reynolds number airfoil were examined. Aerodynamic performance was evaluated through direct measurements of lift, drag, and 1/4-chord pitching moment. Spanwise static pressure and surface film visualization data were also acquired. A reduction in the down-stream airfoil drag coefficient and an increase in its lift/drag were noted in the presence of the canard for a wide range of configurations

Aerodynamic characteristics of the HL-20 and HL-20A lifting-body configurations

Abstract: The data show that the HL-20 is longitudinally and laterally stable over the test range from Mach 10 to 0.2. At hypersonic speeds it has a trimmed lift/drag ratio of 1.4. This values gives the vehicle a cross range capability similar to that of the Space Shuttle. At subsonic speeds, the HL-20 has a trimmed lift/drag ratio of about 3.6. Replacing the flat plate outboard fins with fins having an airfoil shape, increased the maximum trimmed L/D to 4.3. Preliminary evaluation of configuration modifications (the HL-20A series), indicates that trim at higher values of lift at hypersonic speeds could be achieved with an L/D of about 1.0. In the supersonic range, the lift and directional stability characteristics were improved. The untrimmed subsonic L/D was increased to 5.8 with airfoil fins.

Inviscid drag prediction for transonic transport wings using a full-potential method

Abstract: A theory of drag analysis in full-potential flow, based on generalization and extension of Garabedian's and McFadden's idea of determining wave drag by volume integration of the artificial viscosity, is summarized. The applicability of the theory is restricted to shock-capturi ng numerical methods (such as, e.g., finite-volume methods), where artificial viscosity is essential for proper performance. Two mesh refinement experiments on nested grids have been carried out for the DFVLR-F4-wing in transonic flow, using CH- as well as CO-topology grids. The MATRICS code used in the experiments is first-order accurate in the mesh size throughout supersonic flow regions. It is concluded that CO-topology grids are better suited for drag analysis than are CH-topology grids. It is also concluded that the accuracy of each individual drag component can be improved by extrapolating to the limit of vanishing mesh size. Finally, to avoid excessively fine grids in an engineering environment, the need is stressed for artificial viscosity terms that are second-order small in the mesh size in supersonic flow regions, except for the immediate vicinity of the shock waves. However, extrapolations procedures are believed to remain necessary for accurate drag prediction.

Determination of balloon drag

Abstract: The evolution of an empirical drag relationship that has stimulated rethinking regarding the physics of balloon drag phenomena is discussed. Combined parasitic drag from all sources in the balloon system are estimated to constitute less than 10 percent of the total system drag. It is shown that the difference between flight-determined drag coefficients and those based on the spherical assumption should be related to the square of the Froude number.

Waves and thermodynamics in high Mach number propulsive ducts

Abstract: The propulsive effects of waves in ducts, especially at high Mach numbers, are investigated, focusing on drag and thrust and on the conversion of heat into waves which produce thrust. It is shown that essentially all of the work done by an expanding fluid passing through a duct at high Mach number is delivered in the form of waves, and that duct surface angles exist that are optimum for the production of thrust from a wave. The effects of wave phenomena on drag and thrust are considered by extending the concept of a Busemann biplane into that of a 'Busemann scramjet, taking 'off-design' performance into account. An idealized model of a streamtube with heat addition is developed, and flow mechanisms involved in generating thrust by the expansion of this streamtube in an exhaust nozzle are examined.

Aerodynamic shape optimization of arbitrary hypersonic vehicles

Abstract: A new method was developed to optimize, in terms of aerodynamic wave drag minimization, arbitrary (nonaxisymmetric) hypersonic vehicles in modified Newtonian flow, while maintaining the initial volume and length of the vehicle. This new method uses either a surface fitted Fourier series to represent the vehicle's geometry or an independent point motion algorithm. In either case, the coefficients of the Fourier series or the spatial locations of the points defining each cross section were varied and a numerical optimization algorithm based on a quasi-Newton gradient search concept was used to determine the new optimal configuration. Results indicate a significant decrease in aerodynamic wave drag for simple and complex geometries at relatively low CPU costs. In the case of a cone, the results agreed well with known analytical optimum ogive shapes. The procedure is capable of accepting more complex flow field analysis codes.

Effects of turbulence and number density on the drag coefficient of droplets

Abstract: Droplet drag coefficients for polydispersed drops are determined via the behavior of drops decelerating on the stagnation streamline of a cylinder with an afterbody mounted in a wind tunnel test section. A variety of velocity, turbulence levels, and droplet number densities were studied. A force balance equation technique was used to determine drag coefficient. For the levels of number density, up to 700/cc, and turbulence, up to about 7 percent, no definite effects were seen. However, the smallest drops in the high turbulence case showed some evidence of drop-turbulence and/or drop-drop interactions. The drag results that were developed for this set of measurements agreed well with other empirical relations previously determined.

Analysis of Fluctuating Static Pressure Measurements in a Large High Reynolds Number Transonic Cryogenic Wind Tunnel

Abstract: Dynamic measurements of fluctuating static pressure levels were made using flush mounted high frequency response pressure transducers at eleven locations in the circuit of the National Transonic Facility (NTF) over the complete operating range of this wind tunnel. Measurements were made at test section Mach numbers from 0.2 to 1.2, at pressure from 1 to 8.6 atmospheres and at temperatures from ambient to -250 F, resulting in dynamic flow disturbance measurements at the highest Reynolds numbers available in a transonic ground test facility. Tests were also made independently at variable Mach number, variable Reynolds number, and variable drivepower, each time keeping the other two variables constant thus allowing for the first time, a distinct separation of these three important variables. A description of the NTF emphasizing its flow quality features, details on the calibration of the instrumentation, results of measurements with the test section slots covered, downstream choke, effects of liquid nitrogen injection and gaseous nitrogen venting, comparisons between air and nitrogen, isolation of the effects of Mach number, Reynolds number, and fan drive power, and identification of the sources of significant flow disturbances is included. The results indicate that primary sources of flow disturbance in the NTF may be edge-tones generated by test section sidewall re-entry flaps and the venting of nitrogen gas from the return leg of the tunnel circuit between turns 3 and 4 in the cryogenic mode of operation. The tests to isolate the effects of Mach number, Reynolds number, and drive power indicate that Mach number effects predominate. A comparison with other transonic wind tunnels shows that the NTF has low levels of test section fluctuating static pressure especially in the high subsonic Mach number range from 0.7 to 0.9.

Super/hypersonic aerodynamic characteristics for a transatmospheric vehicle concept having a minimum drag forebody

Abstract: Experimental longitudinal and lateral-directional aerodynamic characteristics were obtained for a generic transatmospheric vehicle concept having a replaceable minimum drag forebody shape. The alternate forebody tested was a 1/4-power series body. Tests were made over a range of Mach numbers from 2 to 10 at a nominal Reynolds number, based on a length of 2.3 x 10 to the 8th and angles of attack from -4 to 20 deg. The minimum drag forebody provided significant improvements in minimum drag and L/D for the configuration as well as a longitudinally stabilizing increment. Although the baseline configuration is longitudinally unstable, the L/D improvements at low to moderate angles of attack would enhance cruise performance. Varying wing incidence angles was demonstrated as an effective horizontal trim device without significant trim drag penalties.

NACA0012 Airfoil Data Corrected for Sidewall Boundary-layer Effects in the NAL Two-Dimensional Transonic Wind Tunnel

Abstract: An experimental study on NACA0012 airfoils at transonic speeds has been conducted in order to acquire aerodynamic data for evaluating sidewall boundary-layer effects. Measurements primarily consisted of static pressure on the airfoil and drag force coefficients determined using the wake rake. The tests were performed at free-stream Mach numbers from approximately 0.65 to 0.8, at angles of attack from -2.00 to 2.00, and at Reynolds numbers (based on airfoil chord) from 7*106 to 40 * 106. Mach number corrections for sidewall boundary-layer effects were made, and tests were subsequently performed on two different chord airfoils to confirm the applicability of the correction to different airfoil aspect ratios. This is a supplementary report and presents both uncorrected and corrected data which enables comparative studies to be conducted using computational fluid dynamics and other wind tunnel experimental results.

Evaluation of the drag force by integrating the energy dissipation rate in stokes flow for 2D domains using the FEM

Abstract: SUMMARY The total drag force on the surface of a body, which is the sum of the form drag and the skin friction drag in a 2D domain, is numerically evaluated by integrating the energy dissipation rate in the whole domain for an incompressible Stokes fluid. The finite element method is used to calculate both the energy dissipation rate in the whole domain as well as the drag on the boundary of the body. The evaluation of the drag and the energy dissipation rate are post-processing operations which are carried out after the velocity field and the pressure field for the flow over a particular profile have been obtained. The results obtained for the flow over three different but constant area profiles -a circle, an ellipse and a cross-section of a prolate spheroid-with uniform inlet velocity are presented and it is shown that the total drag force times the velocity is equal to the total energy dissipation rate in the entire finite flow domain. Hence, by calculating the energy dissipation rate in the domain with unit velocity specified at the far-field boundary enclosing the domain, the drag force on the boundary of the body can be obtained.

Experimental and Theoretical Investigation of Vortex Formation in the Wake of a Flat Plate for Subsonic and Supersonic Free-Stream Mach Number

Abstract: The supersonic turbulent wake can be changed to a vortical wake if the surface of the plate is roughened by pasting sandpaper on to it and if the distance d between the trailing edge and the point where the sandpaper begins is in the order of the thickness of the plate II. For d & II the wake contains the strongest vortices while enlarging or diminishing this distance leads to a gradual disappearance of the vortices. If the roughness is increased the strength of the vortices is increased too.It turns out that the decisive parameter which determine the development of the wake are the values of the gradient du/dy of the most inner part of the boundary layer at the trailing edge and its momentum thickness. These two parameters play an important role in the linear stability analysis of wake flows too. A linear local stability analysis with measured velocity profiles of the near wake for subsonic freestream Mach numbers was carried out. It results in a good correspondence between measured and calculated Strouhal numbers. The underlying physical model of the analysis is able to explain qualitatively some aspects of the phenomenon. The rouhgness and the distance d seem to be necessary to reach the required length of the locally absolutely unstable region directly behind the flat plate. The global resonance criterion from Koch can be applied to supercritical Reynolds numbers.

On the process of Mach wave generation in air

Abstract: The process of Mach wave generation in air is studied in both plane and spherical geometries The experimental results reported here are theoretically interpreted using the predictions of a self-similar model of strong explosion along with the hydrodynamic equations of a perfect gas, and a good agreement is found

Aerodynamics of spheres for Mach numbers from 0.6 to 10.5 including some effects of test conditions

Abstract: Wind tunnel tests were made for spheres of various sizes over a range of Mach numbers and Reynolds numbers. The results indicated some conditions where the drag was affected by changes in the afterbody pressure due to a shock reflection from the tunnel wall. This effect disappeared when the Mach number was increased for a given sphere size or when the sphere size was decreased for a given Mach number. Drag measurements and Schlieren photographs show the possibility of obtaining inaccurate data when tests are made with a sphere too large for the test section size and Mach number. Tests were also made of an oblate spheroid. The results indicated a region at high Mach numbers where inherent positive static stability might occur with the oblate-face forward. The drag results are compared with those for a sphere and those for various other shapes. The drag results for the oblate spheriod and the sphere are also compared with some calculated results.

The Near-Field Flow and Drag on Cylindrical Bodies Moving Concentrically Inside Very Long Tubes.

Abstract: : A differential simulation based on the k-model of turbulence describing the flow field around a streamlined body traveling through f1uid along the centerline of a closed end tube has been used to predict drag coefficients for a range of Reynolds numbers and diameter ratios. The range of interest corresponds to torpedo/tube combinations of interest to the U.S. Navy. Pressure coefficients are also plotted as a function of axial position along the body. A finite difference solution of the inviscid flow field is also developed and presented. Comparison of inviscid pressure coefficients with viscous pressure coefficients reveals that the nose region displays essentially inviscid behavior. The viscous differential model verified the hypothesis that total drag on the body could be found by independent calculation of nose drag, cylindrical section drag and wake drag, proving that nose drag and tail drag are independent of the length of the cylindrical section. A one dimensional control volume analysis was performed to predict drag coefficients as a function of Reynolds number.

Euler code evaluation of a transatmospheric vehicle at supersonic speeds

Abstract: An Euler code study has been conducted on a transatmospheric vehicle at Mach numbers from 1.5-4.5. The results show that the Euler code accurately predicted the lift, drag, and pitching-moment coefficients as a function of Mach number and angle of attack. The longitudinal stability characteristics were also predicted reasonably well as a function of Mach number. The code also predicted the changes in lift, drag, and pitching moment for forward and rearward shift of the wing from the nominal position. The Euler code provided an accurate prediction of the aerodynamic characteristics across the Mach-number range.

Mach number validation of a new zonal CFD method (ZAP2D) for airfoil simulations

Abstract: A closed-loop overlapped velocity coupling procedure has been utilized to combine a two-dimensional potential-flow panel code and a Navier-Stokes code. The fully coupled two-zone code (ZAP2D) has been used to compute the flow past a NACA 0012 airfoil at Mach numbers ranging from 0.3 to 0.84 near the two-dimensional airfoil C(lmax) point for a Reynolds number of 3 million. For these cases, the grid domain size can be reduced to 3 chord lengths with less than 3-percent loss in accuracy for freestream Mach numbers through 0.8. Earlier validation work with ZAP2D has demonstrated a reduction in the required Navier-Stokes computation time by a factor of 4 for subsonic Mach numbers. For this more challenging condition of high lift and Mach number, the saving in CPU time is reduced to a factor of 2.

Winglet effectiveness on low aspect ratio wings at supersonic mach numbers

Abstract: A computational study has been conducted on two wings of aspect ratios 1.244 and 1.865, each having 65-deg leading edge sweep angles, to determine the effects of nonplanar winglets at supersonic Mach numbers. A design Mach number of 1.62 was selected. The winglets studied were parametrically varied in alignment, length, sweep, camber, and thickness to determine the effects of winglet geometry on predicted performance. For the computational analysis, an existing Euler code that employed a marching technique was used. The results indicated that the possibility existed for wing-winglet geometries to equal the performance of wing-alone bodies in supersonic flows with both bodies having the same semispan length. The performance parameters of main interest were the lift-to-pressure drag ratio and the pressure drag coefficient as functions of lift coefficient. The lift coefficient range for this study was from -0.20 to 0.70 with emphasis on the range of 0.10 to 0.22.