Showing papers on "Pitching moment published in 1991"

Unsteady flow past an airfoil pitching at a constant rate

Abstract: The unsteady flow past a NACA 0012 airfoil that is undertaking a constant-rate pitching up motion is investigated experimentally by the PIDV technique in a water towing tank. The Reynolds number is 5000, based upon the airfoil's chord and the free-stream velocity. The airfoil is pitching impulsively from 0 to 30 deg. with a dimensionless pitch rate alpha of 0.131. Instantaneous velocity and associated vorticity data have been acquired over the entire flow field. The primary vortex dominates the flow behavior after it separates from the leading edge of the airfoil. Complete stall emerges after this vortex detaches from the airfoil and triggers the shedding of a counter-rotating vortex near the trailing edge. A parallel computational study using the discrete vortex, random walk approximation has also been conducted. In general, the computational results agree very well with the experiment.

Fully coupled implicit method for thermochemical nonequilibrium air at suborbital flight speeds

Abstract: A CFD technique is described in which the finite-rate chemistry in thermal and chemical nonequilibrium air is fully and implicitly coupled with the fluid motion. Developed for use in the suborbital hypersonic flight speed range, the method accounts for nonequilibrium vibrational and electronic excitation and dissociation, but not ionization. The steady-state solution to the resulting system of equations is obtained by using a lower-upper factorization and symmetric Gauss-Seidel sweeping technique through Newton iteration. Inversion of the left-hand-side matrices is replaced by scalar multiplications through the use of the diagonal dominance algorithm. The code, named CENS2H (Compressible-Euler-Navier-Stokes Two-Dimensional Hypersonic), is fully vectorized and requires about 8.8 x 10 to the -5th sec per node point per iteration using a Cray X-MP computer. Converged solutions are obtained after about 2400 iterations. Sample calculations are made for a circular cylinder and a 10 percent airfoil at 5 deg angle of attack. The calculated cylinder flow field agrees with that obtained experimentally. The code predicts a 10 percent change in lift, drag, and pitching moment for the airfoil due to the thermochemical phenomena.

Interference-free measurements of the subsonic aerodynamics of slanted-base ogive cylinders

Abstract: Drag, lift, pitching moment, and base-pressure measurements have been made, free of support interference, on a range of slanted-base ogive cylinders, using the NASA Langley Research Center 13-in magnetic suspension and balance system. Test Mach numbers were in the range 0.04-0.2. Two types of wake flow were observed, a quasi-symmetric turbulent closure or a longitudinal vortex flow. Aerodynamic characteristics differ dramatically between the two wake types. Drag measurements are shown to be in agreement with previous tests. A hysteretic behavior of the wake with varying Reynold's number has been discovered for the 45-deg base. An interaction between forebody boundary-layer state and wake flow and base pressures has been detected for higher slant angles.

Generalized multi-point inverse airfoil design

Abstract: In a rather general sense, inverse airfoil design can be taken to mean the problem of specifying a desired set of airfoil characteristics, such as the airfoil maximum thickness ratio, pitching moment, part of the velocity distribution or boundary-layer development, etc., then from this information determine the corresponding airfoil shape. This paper presents a method which approaches the design problem from this perspective. In particular, the airfoil is divided into segments along which, together with the design conditions, either the velocity distribution or boundary-layer development may be prescribed. In addition to these local desired distributions, single parameters like the airfoil thickness can be specified. The problem of finding the airfoil shape is determined by coupling an incompressible, inviscid, inverse airfoil design method with a direct integral boundary-layer analysis method and solving the resulting nonlinear equations via a multidimensional Newton iteration technique. The approach is fast and easily allows for interactive design. It is also flexible and could be adapted to solving compressible, inverse airfoil design problems.

Dynamic stall on wind turbine blades

Abstract: Dynamic loads must be predicted accurately in order to estimate the fatigue life of wind turbines operating in turbulent environments. Dynamic stall contributes to increased dynamic loads during normal operation of all types of horizontal-axis wind turbine (HAWTs). This report illustrates how dynamic stall varies throughout the blade span of a 10 m HAWT during yawed and unyawed operating conditions. Lift, drag, and pitching moment coefficients during dynamics stall are discussed. Resulting dynamic loads are presented, and the effects of dynamic stall on yaw loads are demonstrated using a yaw loads dynamic analysis (YAWDYN). 12 refs., 22 figs., 1 tab.

Aerodynamics of airfoils with vortex trapped by two spanwise fences

Abstract: An airfoil configuration is presented whose lift is enhanced by a trapped-vortex flow field. Based on the research to date, it is recommended that two spanwise fences be used to enclose the trapped vortex and that the fence heights be adjusted so that the equilibrium condition can be achieved with little or no mass removal from the core region of the vortex. It is also shown that the vortex bubble can be located fore and aft on the airfoil to control aerodynamic parameters, such as the pitching moment. Applications of the high-lift concept presented here are briefly discussed.

Transonic Airfoil Design by Constrained Optimization

Abstract: A N aerodynamic design method is developed which cou- ples flow analysis and numerical optimization to find an airfoil shape with improved aerodynamic performance. The flow analysis code is based on the coupled Euler and bound- ary-layer equations in order to include the rotational, viscous physics of transonic flows. The numerical optimization pro- cess searches for the best feasible design for the specified design objective and design constraints. The method is dem- onstrated with several examples at transonic flow conditions. Contents The optimization process is performed with a commercially available constrained optimization tool.3 The sensitivity of the flow to the perturbation is calculated by finite differences. The effectiveness and efficiency of the design process are influenced by many factors: the number and the shape of the base functions, the number and the tolerance of the con- straints, the flow model and the grid used for flow analyses, and the flight condition at the design point. Design Demonstration The objective of the present design is to produce minimum drag at a specified transonic flight condition. Inequality con- straints are imposed on lift, pitching moment, and cross-sec- tional area of the optimized airfoil. The lift and the area of the optimized airfoil should not be smaller than those of the original airfoil, and the pitching moment should not increase in absolute value. Also imposed are side constraints which limit the magnitude of the design variables. Side constraints are important because a large geometry change can cause boundary-layer separation leading to a termination of the flow solver.

Transonic Navier-Stokes Computations for a Spinning Body of Revolution

Abstract: : A zonal, implicit, time-marching Navier-Stokes computational technique has been used to compute three dimensional transonic flow fields over a projectile. Flow field computations have been performed at M = 0.94 for spin rates of 0 and 4900 rpm and at angles of attack, alpha = 0,4, and 10 degrees. All the computations have been performed on the Cray-2 supercomputer. Details of the flow field such as Mach number contours and surface pressure distributions are presented. Computer surface pressures are compared with available experimental data for the same conditions and the same configuration. Computer results show the large circumferential pressure distribution over the boattail region as well as the nonlinear effect of the angle of attack. Aerodynamic force and moment coefficients (normal force, pitching moment, Magnus force, and Magnus moment) have been obtained from the computed pressures and are compared with the data. The computed results are generally in good agreement with the data at low angle of attack for both nonspinning and spinning conditions. At high angle of attack the agreement is good for the nonspinning case and is less satisfactory for the spinning case.

An improved three-dimensional aerodynamics model for helicopter airloads prediction

Abstract: The importance of unsteady aerodynamic effects due to rotor blade motion is demonstrated for a helicopter in high-speed forward flight. These unsteady effects are modeled as surface boundary conditions to a three-dimensional, unsteady Computational Fluid Dynamics (CFD) code called the Full-Potential Rotor Code (FPR). These boundary conditions cause significant changes in the computed lift and pitching moment at the front and rear of the rotor disk. Airloads from the modified FPR code are then iteratively coupled with the helicopter comprehensive code, CAMRAD/JA. Computed airloads show good agreement with flight-test data when lift values from the FPR code are used in the coupled calculation. However, the computed airloads from CAMRAD/JA along also show good agreement with the experimental data. Thus for this case one cannot demonstrate a significant improvement in computed airloads with the hybrid coupled scheme. The addition of the pitching moment values from the FPR code into the CAMRAD/JA calculation slows down the overall iterative convergence and does not yield any improvement in the final results.

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.

Research and development of high-performance airfoil sections for vertical axis wind turbine at low-Reynolds number.

Abstract: This paper describes a study on the development of a proper airfoil section for the vertical axis wind turbine (VAWT). We aim to abtain pitching moment in the negative direction without losing the excellent characteristics of a symmetrical-type airfoil section. Thus, we propose an increase of the thickness of the symmetrical-type airfoil by giving a meanline reversal, that is, negative meanline near leading edge and positive one near trailing edge, measured from blade chord line. Thus, the airfoil section can be designed by setting the parameters of a reverse meanline and basic airfoil section, and this airfoil section will be called the TW series airfoil section. From the three component force characteristics measured in wind tunnel experiment, TW series airfoil section is regarded to preferable for VAWT, thus, validity of design concept of TW series airfoil section using reverse meanline was demonstrated.

Effect of a simulated ice accretion on the aerodynamics of a swept wing

Abstract: The effect of a simulated glaze ice accretion on the aerodynamic performance of a three-dimensional swept wing is studied experimentally. A semispan wing of effective aspect ratio four was mounted from the sidewall of the UIUC subsonic wind tunnel. The model uses a NACA 0012 airfoil section on a rectangular planform with interchangeable tip and root sections to allow for 0- and 30-degree sweep. A sidewall suction system is used to minmize the tunnel boundary-layer interaction with the model. Surface pressure data from five spanwise stations are compared to earlier data from a similar tunnel. A three-component sidewall balance has been designed, built and used to measure lift, drag and pitching moment on the clean and iced model. The data compare well to the integrated pressure data and to theory on the clean model. In addition, helium-bubble flow visualization has been performed on the iced model and reveals extensive spanwise flow in the separation bubble aft of the upper surface horn. This compares well to the computational results of other researchers. Sidewall suction was found to have no effect on the aerodynamics of the swept wing.

Navier-Stokes simulation of a close-coupled canard-wing-body configuration

Abstract: The thin-layer Navier-Stokes equations are solved for the flow about a coplanar close-coupled canard-wing-body configuration at a transonic Mach number of 0.90 and at angles of attack ranging from 0 to 12 degrees. The influence of the canard on the wing flowfield, including canard-wing vortex interaction and wing vortex breakdown, is investigated. A study of canard downwash and canard leading-edge vortex effects, which are the primary mechanisms of the canard-wing interaction, is emphasized. Comparisons between the computations and experimental measurements of surface pressure coefficients, lift, drag and pitching moment data are favorable. A grid refinement study for configurations with and without canard shows that accurate results are obtained using a refined grid for angles of attack where vortex burst is present. At an angle of attack of approximately 12 deg, favorable canard-wing interaction which delays wing vortex breakdown is indicated by the computations and is in good agreement with experimental findings.

Hypersonic flight experience

Abstract: The flight test results from the X-15, Asset, Prime, Reentry F and Shuttle Orbiter flight research programmes are reviewed and compared with theory and ground-based experiments. Primary emphasis is placed on our present capability to predict aerodynamic coefficients and stability derivatives, and distributions of surface pressure and aerodynamic heating rate for typical orbital and sub-orbital hypersonic vehicles. Overall, this comparison demonstrates the feasibility of designing hypersonic vehicles based on tests in conventional perfect-gas wind tunnels supplemented by state-of-the-art CFD analysis. At Mach numbers up to approximately 8, real-gas effects are small and Mach number/Reynolds number simulation is sufficient to insure accurate prediction of aerodynamic characteristics. At higher mach numbers, real-gas effects become important but appear to affect primarily pitching moment and to have little influence on other aerodynamic characteristics. Viscous interaction effects appear to be well correlated by the viscous interaction parameter, $\overline{V}\_{\infty}^{\prime}$ and to affect primarily axial force. With the exception of RCS jet interactions, stability and control derivatives are well predicted throughout the hypersonic flight regime. State-of-the-art aerodynamic heating techniques appear to give accurate predictions for laminar and fully turbulent attached flows so long as there are no strong shock interactions or non-equilibrium chemistry effects. For such flows, heating rate distributions depend only weakly on Mach number for M$\_{\infty}\gtrsim $ 8 and hence M$_{\infty}\approx $ 8-10 wind tunnel results can be used throughout the hypersonic speed range. For high-speed, high-altitude flight, surface catalycity effects can have a major influence on heating rates but modern, finite-rate boundary layer analyses are capable of predicting major trends. The major remaining challenges are the accurate prediction of: (1) real-gas effects on longitudinal trim, (2) the effectiveness of blended high altitude control systems, (3) shock interaction heating, (4) heating rates for separated vortex-dominated leeside flows, and (5) boundary layer transition and relaminarzation. Finally, it is pointed out that there are no flight data on the propulsion-airframe integration effects that are so important for airbreathing launch vehicles.

Navier-Stokes Predictions of Pitch-Damping for a Family of Flared Projectiles

Abstract: : A parabolized Navier-Stokes approach has been applied to predict the pitch-damping force and moment coefficients for a family of flared projectiles. This family of flared projectiles, which has been tested in the Army Research Laboratory's (ARL) Aerodynamic Range, consists of a common forebody with various flared afterbodies including straight flares, biconic flares, straked flares, and flares with finlets. The predicted pitch-damping force and moment coefficients are determined from the aerodynamic side force and moment acting on the projectiles due to steady coning motion. The predictions of the flow field about the projectiles undergoing steady coning motion are accomplished using a rotating coordinate frame which rotates at the coning rate of the projectile. The governing equations have been modified to include the centrifugal and Coriolis force terms due to the rotating coordinate frame. The predictions of the pitch-damping moment coefficient are compared with pitch-damping coefficient determined from the inflight motion of the projectile. The predictions of the pitch-damping moment coefficients are in good agreement with the range data.

Two complementary approaches to estimate downwash lag effects from flight data

Abstract: ALIDATION of the wind-tunnel and analytical estimates of the aerodynamic derivatives with estimates from flight test data is an important application of the system identification methodology. Reliable and accurate estimates of a large number of aerodynamic derivatives are obtained from flight data using the maximum likelihood method, although, routinely, explicit accounting for certain unsteady aerodynamic effects such as downwash and sidewash interferences pose difficulties. Recent advances in both parameter estimation methods and in flight test techniques have provided a new impetus for modeling and identification of such unsteady interference effects in aircraft dynamics.1'3 This Note addresses in some depth the problem of accounting for downwash lag effects in aircraft parameter estimation from flight data. One possibility of including such an effect is to proceed from the fundamental and to model the forces acting on the wing and tail plane separately. In such a case it becomes possible to estimate explicitly the downwash angle and also to include the lag effect. It leads, however, to a model that is nonlinear in parameters and necessarily requires an estimation program capable of handling nonlinear model postulates. The other approach is a simplification based on linearization. It leads to the first-order approximation of aerodynamic derivatives with respect to translational acceleration. The derivatives due to the body-fixed translational acceleration in the vertical direction are equivalent in the stability axis to those with respect to the rate of change of angle of attack. The second approach was first investigated in Ref. 1 to estimate from flight data the pitching moment derivatives with respect to pitch rate and rate of change of angle of attack of a highly maneuver able aircraft with large roll angle capabilities. A parameter estimation program capable of handling linear as well as nonlinear system models is used here to compare the two approaches of accounting for the downwash. Further, an attempt is made to investigate the possibility of using the linearized approach to estimate separately the two pitch damping derivatives from flight tests with a larger aircraft having limited roll angle capabilities.

Unsteady Navier-Stokes computations on a wing-body configuration in ramp motions

Abstract: Unsteady Navier-Stokes computations are conducted for transonic flows over a wing-body configuration undergoing prescribed ramp motions. The ramp motion from 0 to 15 deg includes angles of attack where vortex breakdown is observed experimentally. The vortex breakdown is found to be delayed until after the ramp motion ends. The dynamic effects on the loads are also demonstrated. The moment coefficient is found to be sensitive to the effect of the virtual mass of the fluid. To verify the numerical results, a grid refinement study has been performed for both steady and unsteady flow conditions, by using up to approximately one million grid points. The results show that the coarse grid can give integrated quantities reasonably well, whereas the finer grids give a more detailed flow structure. Comparisons are also made with available steady-state experimental data.

Testing apparatus for resistance of ship

Abstract: PURPOSE:To make it possible to study and try the combinations of design elements in the wide range by supporting a model ship through a three- component meter, and providing an attitude computer, a coordinate converter, a recorder and a feedback control system. CONSTITUTION:A model ship 10 is supported with a tow tractor 2 through a slant-angle changing actuator 5, a floating amount changing actuator 6 and a three-component meter 11 for measuring drag, pitching moment and floating force. The actuators 5 and 6 are attached to the meter 11. An attitude computer 14 computes the attitude of a model ship based on the signals from displacement detectors 12 and 13 which are attached to the actuators 5 and 6. A coordinate converter 15 obtain three components for the fixed coordinates of the tow tractor 2 based on the signals from the three-component meter 11. The measured values of the three-components are recorded in a recorder 16. In a feedback control system FB, the measured values of the pitching moment and the floating force among the measured values are compared with respective preset values. The actuators 5 and 6 are operated so that the deviations become zero. Switches 21 and 22 for ON/OFF operations are provided in the control system FB.

Local and overall aerodynamic coefficients for bodies in hypersonic, rarefied flow

Abstract: A computational method is given for the prediction of local pressure and viscous shear stress on windward surfaces of bluff, convex, axisymmetric or quasi-axisymmetric, hypersonic bodies in the transitional, rarefied flow regime. Overall aerodynamic forces and moments are computed by integration of the local quantities. The method is based upon a correlation of local pressure and shear stress computed by the direct simulation Monte Carlo (DSMC) numerical technique for cold wall, real gas conditions and some supplemental data from low-density, hypersonic wind tunnels. The relative simplicity of the method makes it feasible to do the necessary calculations with a personal computer. Two-dimensional shapes and leeward surfaces are not included in the scope of the method as it is presented here. Results are compared with DSMC computations for both local and overall coefficients. The latter includes sphere and blunt cone drag as well as lift and pitching moment coefficients for the NASA AFE vehicle at various angles of attack. Very satisfactory agreement is shown.

Effect of exhaust plume/afterbody interaction on installed scramjet performance

Abstract: Newly emerging aerospace technology points to the feasibility of sustained hypersonic flight. Designing a propulsion system capable of generating the necessary thrust is now the major obstacle. First-generation vehicles will be driven by air-breathing scramjet (supersonic combustion ramjet) engines. Because of engine size limitations, the exhaust gas leaving the nozzle will be highly underexpanded. Consequently, a significant amount of thrust and lift can be extracted by allowing the exhaust gases to expand along the underbody of the vehicle. Predicting how these forces influence overall vehicle thrust, lift, and moment is essential to a successful design. This work represents an important first step toward that objective. The UWIN code, an upwind, implicit Navier-Stokes computer program, has been applied to hypersonic exhaust plume/afterbody flow fields. The capability to solve entire vehicle geometries at hypersonic speeds, including an interacting exhaust plume, has been demonstrated for the first time. Comparison of the numerical results with available experimental data shows good agreement in all cases investigated. For moderately underexpanded jets, afterbody forces were found to vary linearly with the nozzle exit pressure, and increasing the exit pressure produced additional nose-down pitching moment. Coupling a species continuity equation to the UWIN code enabled calculations indicating that exhaust gases with low isentropic exponents (gamma) contribute larger afterbody forces than high-gamma exhaust gases. Moderately underexpanded jets, which remain attached to unswept afterbodies, underwent streamwise separation on upswept afterbodies. Highly underexpanded jets produced altogether different flow patterns, however. The highly underexpanded jet creates a strong plume shock, and the interaction of this shock with the afterbody was found to produce complicated patterns of crossflow separation. Finally, the effect of thrust vectoring on vehicle balance has been shown to alter dramatically the vehicle pitching moment.

Aerodynamic measurements on a finite wing with simulated ice

Abstract: The effect of a simulated glaze ice accretion on the aerodynamic performance of a three-dimensional straight and swept wing is studied experimentally. A semispan wing of effective aspect ratio five was mounted from the sidewall of the UIUC subsonic wind tunnel. The model uses an NACA 0012 airfoil section on a rectangular planform with interchangeable tip and root sections to allow for 0- and 30-deg sweep. A sidewall suction system is used to minimize the tunnel boundary-layer interaction with the model. A three-component sidewall balance has been designed, built and used to measure lift, drag and pitching moment on the clean and iced model. Fluorescent oil flow visualization has been performed on the iced model and reveals extensive spanwise flow in the separation bubble aft of the upper surface horn. These results are compared to computational results for the surface pressures, span loads and surface oil flow.

Validation of engineering methods for predicting hypersonic vehicle controls forces and moments

Abstract: This work examines the ability of the aerodynamic analysis methods contained in an industry standard conceptual design code, the Aerodynamic Preliminary Analysis System (APAS II), to estimate the forces and moments generated through control surface deflections from low subsonic to high hypersonic speeds. Predicted control forces and moments generated by various control effectors are compared with previously published wind-tunnel and flight-test data for three vehicles: the North American X-15, a hypersonic research airplane concept, and the Space Shuttle Orbiter. Qualitative summaries of the results are given for each force and moment coefficient and each control derivative in the various speed ranges. Results show that all predictions of longitudinal stability and control derivatives are acceptable for use at the conceptual design stage.

On the estimation of jet-induced fountain lift and additional suckdown in hover for two-jet configurations

Abstract: Currently available methods for estimating the net suckdown induced on jet V/STOL aircraft hovering in ground effect are based on a correlation of available force data and are, therefore, limited to configurations similar to those in the data base. Experience with some of these configurations has shown that both the fountain lift and additional suckdown are overestimated but these effects cancel each other for configurations within the data base. For other configurations, these effects may not cancel and the net suckdown could be grossly overestimated or underestimated. Also, present methods do not include the prediction of the pitching moments associated with the suckdown induced in ground effect. An attempt to develop a more logically based method for estimating the fountain lift and suckdown based on the jet-induced pressures is initiated. The analysis is based primarily on the data from a related family of three two-jet configurations (all using the same jet spacing) and limited data from two other two-jet configurations. The current status of the method, which includes expressions for estimating the maximum pressure induced in the fountain regions, and the sizes of the fountain and suckdown regions is presented. Correlating factors are developed to be used with these areas and pressures to estimate the fountain lift, the suckdown, and the related pitching moment increments.

Performance of an aerospace plane propulsion nozzle

Abstract: A novel inviscid and viscous analysis is provided for a nozzle that is used with a scramjet for thrust generation. The analysis is based on the theory of a two-dimensional minimum length nozzle with a curved inlet surface, where the flow is sonic or supersonic. Inlet conditions are prescribed and the gas is assumed to be perfect. Viscous and inviscid nondimensional parametric results are provided for the thrust, lift, heat transfer, pitching moment, and a variety of boundary-laye r thicknesses. In addition to global results, wall distributions of thrust, heat transfer, etc., are provided. The analysis demonstrates that the nozzle produces a lift force whose magnitude may exceed the thrust and a significant pitching moment. The thrust is sensitive to the inlet Mach number; it rapidly decreases as this Mach number increases. There is little loss in the thrust as the nozzle's downstream wall is truncated while the corresponding decrease in lift and the pitching moment is moderate.

Fourier functional analysis for unsteady aerodynamic modeling

Abstract: A method based on Fourier analysis is developed to analyze the force and moment data obtained in large amplitude forced oscillation tests at high angles of attack. The aerodynamic models for normal force, lift, drag, and pitching moment coefficients are built up from a set of aerodynamic responses to harmonic motions at different frequencies. Based on the aerodynamic models of harmonic data, the indicial responses are formed. The final expressions for the models involve time integrals of the indicial type advocated by Tobak and Schiff. Results from linear two- and three-dimensional unsteady aerodynamic theories as well as test data for a 70-degree delta wing are used to verify the models. It is shown that the present modeling method is accurate in producing the aerodynamic responses to harmonic motions and the ramp type motions. The model also produces correct trend for a 70-degree delta wing in harmonic motion with different mean angles-of-attack. However, the current model cannot be used to extrapolate data to higher angles-of-attack than that of the harmonic motions which form the aerodynamic model. For linear ramp motions, a special method is used to calculate the corresponding frequency and phase angle at a given time. The calculated results from modeling show a higher lift peak for linear ramp motion than for harmonic ramp motion. The current model also shows reasonably good results for the lift responses at different angles of attack.

Aerodynamic Coefficients of the M483A1 Determined from Spark Range Tests

Abstract: : The aerodynamic coefficients of the M483A1 have been refined based on analysis of 65 BRL Transonic Range tests. Both the original M483A1 tests (1975) and the latest tests from 1987-1989 were analyzed utilizing a six degree-of- freedom technique during September-October 1990. The Magnus moment was found to be extremely nonlinear with angle of attack and Mach number. The size of the slow arm limit cycle was computed along the region of dynamic instability due to a large positive Magnus moment at moderate yaw levels. The values of the axial force, normal force, pitching moment, damping moment, and Magnus moment are presented.

Hub and Pylon Fairing Integration for Helicopter Drag Reduction

Abstract: The results of testing hub and pylon fairings mounted on a one-fifth scale helicopter with the goal of reducing parasite drag are presented. Lift, drag, and pitching moment, as well as side force and yawing moment, were measured. The primary objective of the test was to validate the drag reduction capability of integrated hub and pylon configurations in the aerodynamic environment produced by a rotating hub in forward flight. In addition to the baseline helicopter without fairings, three hub fairings and three pylon fairings were tested in various combinations. The three hub fairings tested reflect two different conceptual design approaches to implementing an integrated fairing configuration on an actual aircraft. The design philosophy is discussed in detail and comparisons are made between the wind tunnel models and potential full-scale prototypes. The data show that model drag can be reduced by as much as 20.8 percent by combining a small hub fairing with circular arc upper and flat lower surfaces and a nontapered 34-percent thick pylon fairing. Aerodynamic effects caused by the fairings, which may have a significant impact on static longitudinal and directional stability, were observed. The results support previous research which showed that the greatest reduction in model drag is achieved if the hub and pylon fairings are integrated with minimum gap between the two.

Effect of Nonuniform Entrance Flow Profile on Hypersonic Nozzle Pitching Moment

Abstract: : The Scramjet Hypersonic Nozzle (SCHNOZ) parabolized Navier-Stokes computer code was used to model turbulent, chemically reacting flow present in a hypersonic nozzle. Two nozzle configurations were considered, an isolated nozzle (no external flow) and a nozzle with a finite length cowl. A single nonuniform entrance flow profile was generated and an equivalent uniform flow profile calculated for input into the nozzle code. Uniform and nonuniform cases for each nozzle were run using both frozen and finite rate chemistry. An increased grid resolution in the computer code was necessary to eliminate numerically induced anomalies in the results of the nonuniform cases for both nozzle configurations. Comparisons between the finite rate and frozen flow cases showed that chemistry was essentially frozen for the finite rate cases, indicating that for the nozzle inlet conditions and geometry used in this study, the extra computational time spent on finite rate kinetics was unnecessary. The effects of the nonuniform flowfield used in this study included an increase in the overall vehicles thrust and a decrease in the overall vehicle moment.

Prediction techniques for jet-induced effects in hover on STOVL aircraft

Abstract: Prediction techniques for jet induced lift effects during hover are available, relatively easy to use, and produce adequate results for preliminary design work. Although deficiencies of the current method were found, it is still currently the best way to estimate jet induced lift effects short of using computational fluid dynamics. Its use is summarized. The new summarized method, represents the first step toward the use of surface pressure data in an empirical method, as opposed to just balance data in the current method, for calculating jet induced effects. Although the new method is currently limited to flat plate configurations having two circular jets of equal thrust, it has the potential of more accurately predicting jet induced effects including a means for estimating the pitching moment in hover. As this method was developed from a very limited amount of data, broader applications of the method require the inclusion of new data on additional configurations. However, within this small data base, the new method does a better job in predicting jet induced effects in hover than the current method.