Showing papers on "Pitching moment published in 1975"

Augmentation of Vortex Lift by Spanwise Blowing

Abstract: An investigation has been conducted to evaluate the aerodynamic effects associated with blowing a jet spanwise over a wing's upper surface in a direction parallel to the leading edge. Experimental pressure and force data were obtained on wings with sweep angles of 30 and 45 degrees and showed that spanwise blowing aids in the formation and control of the leading-edge vortex and, hence, significantly improves the aerodynamic characteristics at high angles of attack. Full vortex section lift is achieved at the inboard span station with a small blowing rate, but successively higher blowing rates are necessary to attain the full vortex-lift level at increased span distances. Spanwise blowing generates large increases in lift at high angles of attack, improves the drag polars, and extends the linear pitching moment to high lifts.

Design of Low-Speed Airfoils by Numerical Optimization

Abstract: A practical procedure for the optimum design of low-speed airfoils is demonstrated. The procedure uses an optimization program based on a gradient algorithm coupled with an aerodynamic analysis program that uses a relaxation solution of the inviscid, full-potential equation. The analysis program is valid for both incompressible and compressible flow, thereby making optimum design of high-speed, shock-free airfoils possible. Results are presented for the following three constrained optimization problems at fixed angle of attack and Mach number: (1) adverse pressure-gradient minimization, (2) pitching-moment minimization; and (3) lift maximization. All three optimization problems were studied with various aerodynamic and geometric constraints.

The Thin-Hydrofoil Thickness Problem Including Leading-Edge Corrections

Abstract: The method of Keldysh and Lavrentiev is used to study the potential flow of a uniform stream past a submerged thin symmetric hydrofoil. An expression for the surface speed is obtained and the method of matched asymptotic expansions is used to develop corrections at the round leading edge. The pressure, lift, drag, and pitching moment are presented for the Joukowski hydrofoil with emphasis on the variation with thickness, Froude number, and chord-to-depth ratio.

Application of numerical optimization to the design of low speed airfoils

Abstract: A practical procedure for the optimum design of low-speed airfoils is demonstrated. The procedure uses an optimization program based on the method of feasible directions coupled with an aerodynamic analysis program that uses a relaxation solution of the inviscid, full potential equation. Results are presented for airfoils designed to have small adverse pressure gradients, high maximum lift, and low pitching moment.

Aerodynamic characteristics of the 10-percent-thick NASA supercritical airfoil 33 designed for a normal-force coefficient of 0.7

Abstract: This report documents the experimental aerodynamic characteristics of a 14 percent thick supercritical airfoil based on an off design sonic pressure plateau criterion. The design normal force coefficient was 0.7. The results are compared with those of the family related 10 percent thick supercritical airfoil 33. Comparisons are also made between experimental and theoretical characteristics and composite drag rise characteristics derived for a full scale Reynolds number of 40 million.

Shear flow aerodynamics - Lifting surface theory

Abstract: A lifting surface theory based on a parallel shear flow model is presented for steady, incompressible flows. The theory is intended to account approximately for the presence of a boundary layer. The method of Fourier transforms is used to calculate the pressure on a surface of infinite extent and arbitrary contour. Immediately above the surface is a region of sheared flow (the boundary layer), outside of which the flow velocity is constant. The Fourier transform of the pressure on this surface is used to derive the shear flow equivalent to the kernel function of classical potential flow lifting surface theory. The kernel function provides an integral relation between the upwash at a given point on the surface and the pressure everywhere on the surface. This relation is treated as an integral equation for the pressure, and is solved numerically. Computations are presented for the lift and pitching moment on a flat plate in two-dimensional flow, and for flat, rectangular wings of aspect ratio 1, 2, and 5. As expected, the shear layer decreases the lift curve slope; however, the shear layer (whose thickness is constant along the wing chord) has little effect on the center of pressure.

Effect of compressibility on the nonlinear prediction of the aerodynamic loads on lifting surfaces

Abstract: The vortex-lattice technique for incompressible flow which accounts for separation at sharp edges is modified to account for compressibility. This is accomplished by extending the Prandtl-Glauert transformation to moderate angles of attack. Thus, the aerodynamic characteristics for the compressible case are obtained from the solution of an equivalent incompressible problem. Numerical results are presented for parallelogram and delta wings to assess the effects of compressibility. The results are in good agreement with available experimental data.

Effects of forward contour modification on the aerodynamic characteristics of the NACA 641-212 airfoil section

Abstract: Two different forward contour modifications designed to increase the maximum lift coefficient of the NACA 64 sub 1-212 airfoil section were evaluated experimentally at low speeds. One modification consisted of a slight droop of the leading edge with an increased leading-edge radius; the other modification incorporated increased thickness over the forward 35 percent of the upper surface of the profile. Both modified airfoil sections were found to provide substantially higher maximum lift coefficients than the 64 sub 1-212 section. The drooped leading-edge modification incurred a drag penalty of approximately 10 percent at low and moderate lift coefficients and exhibited a greater nosedown pitching moment than the 64 sub 1-212 profile. The upper surface modification produced about the same drag level as the 64 sub 1-212 section at low and moderate lift coefficients and less nosedown pitching moment than the 64 sub 1-212 profile. Both modified airfoil sections had lower drag coefficients than the 64 sub 1-212 section at high lift coefficients.

Static Aerodynamics of Missile Configurations for Mach O to 3

Abstract: Several theoretical and empirical procedures are combined to form a useful design tool for computing static aerodynamics on gun-launched guided projectiles and missiles. The Mach number and angle-of-attack range over which the method is applicable are O^A/^ <3 and 0 < oc < 15°, respectively. Body and wing geometries can be quite general in that pointed or blunt nose bodies and sharp or blunt leading-edge wings can be assumed. Computed results for several configurations compare well with experimental and other analytical results. The computer program is cost effective as it only costs about $5 per Mach number to compute the lift, drag, and pitching moment of a typical wing-body shape on the CDC 6700 Computer.

Jet-Wing Lifting-Surface Theory Using Elementary Vortex Distributions

Abstract: A lifting-surface theory for jet wings based on a finite-element method—the method of elementary vortex distribution or the EV D method—is presented. The method utilizes a set of independent but overlapped elementary horseshoe vortex distributions to represent the wing and jet sheet, and satisfies a set of mixed-type boundary conditions on both the wing and jet sheet. The solution includes chordwise and span wise loading distributions, from which sectional and total aerodynamic quantities (e.g., lift, pitching moment, induced drag, etc.) are derived. In view of the finite-element approach, the method can, in general, be applied to jet wings of arbitrary planform, camber, twist, partial-span flaps, and arbitrary trailing-edge jet-momentum distribution. The present method also reduces to a conventional lifting-surface theory when the jet momentum is zero. An extensive comparison has been made of solutions derived with the EVD method with other theoretical and experimental data for jet wings and conventional wings. Good agreement has been observed in the chordwise and span-wise loadings as well as total aerodynamic coefficients.

An analytical evaluation of airfoil sections for helicopter rotor applications

Abstract: An analytical technique was used to evaluate airfoils for helicopter rotor application. This technique permits assessment of the influences of airfoil geometric variations on drag divergence Mach number at lift coefficients from near zero to near maximum lift. Analytical results presented in this paper indicate the compromises in drag divergence Mach number which result from changes in (1) thickness ratio, (2) location of maximum thickness, (3) leading-edge radius, (4) camber addition, and (5) location of maximum camber of NACA four- and five-digit-series airfoils and some 6-series airfoils of potential interest for helicopters. Examples of airfoil sections which combine several of the geometric changes favorable to both advancing and retreating section performance have been presented.

Experimental and theoretical low speed aerodynamic characteristics of the NACA 65 sub 1-213, alpha equals 0.50, airfoil

Abstract: Low-speed wind-tunnel tests have been conducted to determine the two-dimensional aerodynamic characteristics of the NACA 65 sub 1-213, a = 0.05, airfoil. The results were compared with data from another low-speed wind tunnel and also with theoretical predictions obtained by using a viscous subsonic method. The tests were conducted over a Mach number range from 0.10 to 0.36. Reynolds numbers based on the airfoil chord varied from about 3 million to 23 million.

Calculation of static longitudinal aerodynamic characteristics of STOL aircraft with upper surface blown flaps

Abstract: An existing prediction method developed for EBF aircraft configurations was applied to USB configurations to determine its potential utility in predicting USB aerodynamic characteristics. An existing wing-flap vortex-lattice computer program was modified to handle multiple spanwise flap segments at different flap angles. A potential flow turbofan wake model developed for circular cross-section jets was used to model a rectangular cross-section jet wake by placing a number of circular jets side by side. The calculation procedure was evaluated by comparison of measured and predicted aerodynamic characteristics on a variety of USB configurations. The method is limited to the case where the flow and geometry of the configuration are symmetric about a vertical plane containing the wing root chord. Comparison of predicted and measured lift and pitching moment coefficients were made on swept wings with one and two engines per wing panel, various flap deflection angles, and a range of thrust coefficients. The results indicate satisfactory prediction of lift for flap deflections up to 55 and thrust coefficients less than 2. The applicability of the prediction procedure to USB configurations is evaluated, and specific recommendations for improvements are discussed.

Aerodynamic characteristics of a tandem wing configuration of a Mach number of 0.30

Abstract: An investigation was conducted to determine the aerodynamic characteristics of a tandem wing configuration. The configuration had a low forward mounted sweptback wing and a high rear mounted sweptforward wing jointed at the wing tip by an end plate. The investigation was conducted at a Mach number of 0.30 at angles of attack up to 20 deg. A comparison of the experimentally determined drag due to lift characteristics with theoretical estimates is also included.

Low-Speed Aerodynamic Forces and Moments Acting on the Human Body

Abstract: : Information on the aerodynamic forces acting on the human body is sparse, and scattered in various hard-to-find reports A primary purpose of this report is to collect the available data in one volume, and to present it in the most useful form The aerodynamic force and moment data originally reported by Schmitt for three body positions is reported, together with additional data which he did not publish The drag portion of the Schmitt data is then compared with all other available drag data, represented by wind tunnel tests with volunteer subjects and anthropomorphic dummies, and the instrumented free falls of parachutists and anthropomorphic dummies

A Numerical Design Study of Fully Cavitating Hydrofoil Sections Having Prescribed Pressure Distributions

Abstract: Results of a study of fully cavitating hydrofoil sections are reported. All calculations are based upon the linearized theory of cavity flow in two dimensions. This report is the first detailed exploration of the consequences of the general inverse theory which permits the designer to specify the design values of the lift coefficient, cavitation number and the thickness of the upper surface of the cavity at the profile trailing edge as well as the shape of the pressure distribution on the wetted surface. The ordinates of the upper cavity contour and the wetted surface contour are calculated. The design angle of attack, the cavity length, the drag coefficient and the moment coefficient are also calculated. It is found for almost any cavitation number and any design lift coefficient that if the center of pressure is placed as closely as possible to the profile leading edge the resulting profile will have the most favorable lift-to-drag ratio. The study also includes off- design calculations, in accordance with the direct theory of cavity flows, to determine cavity interference with the upper nonwetted surface of the profile and the hydrodynamic forces of particular designs. (Author)

Experimental Evaluation of the Performance of the TAP-1 Supercavitating Hydrofoil Model at 80 Knots

Abstract: : Experimental results are presented for a model of the TAP-1 supercavitating hydrofoil system in its high-speeds mode of operation. To realistically simulate the ventilation air demand of the prototype craft, the model was examined at full water speeds using cavitation number scaling. The experiments were conducted in the NASA Aircraft Landing Dynamics Facility, an outdoor free-running towing carriage. Unsteady loads in lift, drag, side force, and pitching moment were continuously recorded on analog tape and were then time averaged. The foil (chordline) angle of attack ranged from 2.4 to 10.4 degrees at 80 knots. The strut side force in yaw developed by the foil system and by the basic parabolic strut only was recorded for speeds of 50, 70, and 80 knots. The maximum strut sideslip angle at 80 knots before sudden side ventilation was 3 1/ 4 degrees. The cavity air demand increased linearly with water speed (or Froude number) over the range of speeds examined. While the strut spray wedges were absolutely necessary to achieve full ventilation, their presence added only about 10 percent to the drag. The spanwise twist of the model was successful in maintaining the cavity out to the wing tips at low angles of attack, but failed to recreate the sectional loading. The maximum lift-to-drag ratio measured in full cavity flow was 6.6. No vortex shedding or leading edge vibrations were observed.

Procedures for the design of low-pitching-moment airfoils

Abstract: The pitching moment of a given airfoil is decreased by specifying appropriate modifications to its pressure distribution; by prescribing parameters in a special formula for the Theodorsen epsilon-function; and with superposition of a thickness distribution and subsequent tailoring. Advantages and disadvantages of the three methods are discussed.

Aerodynamic Theory for a Cascade of Oscillating Airfoils in Compressible Subsonic Flow.

Abstract: : A simple numerical technique is developed for determining the airload coefficients on a typical airfoil of the cascade. The effects of varying airfoil spacing, frequency, Mach number and phase difference between adjacent blades are investigated. In particular, the variations in the aerodynamic damping for pure vertical translational and pitching motions are considered. It is shown that translational damping can become zero but never negative at certain discrete frequencies. Furthermore, the results indicate that the pitching moment aerodynamic damping referred to the quarter-chord axis, while also being zero at the critical frequencies, can be negative at the higher Mach numbers over a wide range of frequencies of interest in flutter analysis. The effects of varying the stagger angle of the cascade are not considered in the present paper.

Tunnel Interference Reduction on a Finite Airfoil

Abstract: HE ventilated wind tunnel has been introduced and refined over the last two decades from the fixed geometry to the variable porosity perforated walls. Both theoretical results1 and experimental data2 have demonstrated that it was difficult to eliminate pitching-moment interference of an aircraft model simultaneously with lift interference when using walls with uniformly distributed porosity. The evidence3 in the experimental development of walls for V/STOL testing indicates that it becomes possible to simultaneously eliminate pitching moment and lift interference in a slotted tunnel with nonuniformly distributed porosity. The interference reduction concept has been demonstrated by the author's previous paper4 in which a mathematical technique is presented to predict the interference on an airfoil represented by a single singularity in a Gaussian-type distribution of porosity. The present paper extends the mathematical technique to the case of a finite chord airfoil to predict the proper porosity distribution to eliminate the interference. The present calculation becomes urgent and necessary for an experimental program developing a new "nonadaptive" wall. On the other hand, the self-correcting wind tunnel concept5'6 which refers to the "adaptive" wall, has been demonstrated by a numerical method7 to achieve interference-free flow conditions. This wind tunnel requires the provisions either to adjust the deflections of the wall surfaces, or to adjust the wall porosity and plenum pressure behind the wall to establish unconfined flow conditions. The present result of interference calculation is primarily for the improvement of the current existing nonadaptive wall wind tunnel. Analysis The airfoil is placed at the centerline of a perforated tunnel having walls with nonuniform distribution of porosities. The tunnel interference on a model is divided into two parts in the investigation, i.e. the blockage and lift interferences. Based on the thin airfoil theory, the airfoil thickness and camber may be treated separately. The blockage and lift interferences are induced by the presence of the airfoil thickness and camber in the tunnel, respectively. The formulation for both types of interferences is essentially similar, except for the different mathematical representations of the airfoil thickness and camber. The small perturbation theory for subsonic flow is used in the analysis with a homogeneous boundary conModel in a Tunnel

Low Speed Wind Tunnel Tests on a One-Seventh Scale Model of the H.126 Jet Flap Aircraft

Abstract: Low speed wind tunnel tests were performed on a one-seventh scale model of the British H.126 jet flap research aircraft over a range of jet momentum coefficients. The primary objective was to compare model aerodynamic characteristics with those of the aircraft, with the intent to provide preliminary data needed towards establishing small-to-full scale correlating techniques on jet flap V/STOL aircraft configurations. Lift and drag coefficients from the model and aircraft tests were found to be in reasonable agreement. The pitching moment coefficient and trim condition correlation was poor. A secondary objective was to evaluate a modified thrust nozzle having thrust reversal capability. The results showed there was a considerable loss of lift in the reverse thrust operational mode because of increased nozzle-wing flow interference. A comparison between the model simulated H.126 wing jet efflux and the model uniform pressure distribution wing jet efflux indicated no more than 5% loss in weight flow rate.