Showing papers on "Airfoil published in 1988"

Effects of independent variation of Mach and Reynolds numbers on the low-speed aerodynamic characteristics of the NACA 0012 airfoil section

Abstract: A comprehensive data base is given for the low speed aerodynamic characteristics of the NACA 0012 airfoil section. The Langley low-turbulence pressure tunnel is the facility used to obtain the data. Included in the report are the effects of Mach number and Reynolds number and transition fixing on the aerodynamic characteristics. Presented are also comparisons of some of the results with previously published data and with theoretical estimates. The Mach number varied from 0.05 to 0.36. The Reynolds number, based on model chord, varied from 3 x 10 to the 6th to 12 x 10 to the 6th power.

Validation of approximate indicial aerodynamic functions for two-dimensional subsonic flow

Abstract: Approximations for two-dimensional indicial (step) aerodynamic responses due to angle of attack and pitch rate are obtained and generalized to account for compressibility effects up to a Mach number of 0.8. Using the Laplace transform method, these indicial functions are manipulated to produce explicit solutions for idealized harmonic forcings. These explicit solutions are subsequently compared with experimentally obtained pitch and plunge aerodynamic data in the reduced frequency domain. The results of this comparison are used to relate back and substantiate the generalization of the compressible indicial lift and moment functions.

Boundary-layer measurements on an airfoil at low Reynolds numbers

Abstract: Etude experimentale des caracteristiques des bulles de decollement de transition par anemometrie laser et a fil chaud

Experimental results for the Eppler 387 airfoil at low Reynolds numbers in the Langley low-turbulence pressure tunnel

Abstract: Experimental results were obtained for an Eppler 387 airfoil in the Langley Low Turbulence Pressure Tunnel. The tests were conducted over a Mach number range from 0.03 to 0.13 and a chord Reynolds number range for 60,000 to 460,000. Lift and pitching moment data were obtained from airfoil surface pressure measurements and drag data for wake surveys. Oil flow visualization was used to determine laminar separation and turbulent reattachment locations. Comparisons of these results with data on the Eppler 387 airfoil from two other facilities as well as the Eppler airfoil code are included.

Airfoil dynamic stall at constant pitch rate and high Reynolds number

Abstract: An experiment has been performed to study the aerodynamics of dynamic stall penetration at constant pitch rate and high Reynolds number, in an attempt to model more accurately conditions during aircraft poststall maneuvers and during helicopter high-speed forward flight. An airfoil was oscillated at pitch rates, A = ac/2U between 0.001 and 0.020, Mach numbers between 0.2 and 0.4, and Reynolds numbers between 2-4 x 10. Surface pressures were measured using 72 miniature transducers, and the locations of transition and separation were determined using 8 surface hot-film gages. The results demonstrate the influence of the leading-edge vorticity on the unsteady aerodynamic response during and after stall. The vortex is strengthened by increasing the pitch rate and is weakened by increasing the Mach number and by starting the motion close to the steady-state stall angle. A periodic pressure oscillation occurred after stall at high pitch angle and moderate Reynolds number; the oscillation frequency was close to that predicted for a von Karman vortex street. A small supersonic zone near the leading edge at M = 0.4 was found to reduce significantly the peak suction pressures and the unsteady increments to the airloads. These results provide the first known data base of constant-pitch-rate aerodynamic information at realistic combinations of Reynolds and Mach numbers.

Viscous transonic airfoil workshop compendium of results

Abstract: Results from the Viscous Transonic Airfoil Workshop held at the AIAA 25th Aerospace Sciences Meeting at Reno, NV in January 1987, are compared with each other and with experimental data. Test cases used in this workshop include attached and separated transonic flows for three different airfoils: the NACA 0012 airfoil, the RAE 2822 airfoil, and the Jones airfoil. A total of 23 sets of numerical results from 15 different author groups are included. The numerical methods used vary widely and include: 16 Navier-Stokes methods, 2 Euler/boundary-layer methods, and 5 full-potential/boundary-layer methods. The results indicate a high degree of sophistication among the numerical methods with generally good agreement between the various computed and experimental results for attached or moderately-separated cases. The agreement for cases with larger separation is only fair and suggests additional work is required in this area.

Low-Reynolds-number airfoil design for the M.I.T. Daedalus prototype- A case study

Abstract: The rationale used for the aerodynamic wing design of the prototype long-range human-powered aircraft Light Eagle is presented. Three different airfoils, designed for chord Reynolds numbers of 500,000,375,000, and 250,000 were used across the wingspan. The airfoil design rationale centered on minimizing the losses in the transitional separation bubbles typically occurring on airfoils at Reynolds numbers of less than 1 million. Structural and manufacturing constraints were also a consideration in the airfoil design, although to a lesser extent. Airfoil performance prediction during the design process was done entirely through numerical simulation. The numerical model employs the Euler equations to represent the inviscid flow, and an integral boundary-layer formulation to represent the viscous flow. Strong viscous-inviscid coupling and an amplification transition criterion included in the overall equation system permit calculation of transitional separation bubbles and their associated losses. Flow visualization tests performed on the Light Eagle at various lift coefficients in towed flight revealed transition occurring very near the intended position on the wing surface except within a few chords of the tip, where the flow appeared to be turbulent over most of the upper surface. Total drag aircraft polars obtained from the measured aircraft energy time history in glide contained too much scatter to be used as quantitative test data but did reproduce the basic trends of the calculations, including maximum lift coefficient levels.

Airfoil with nested cooling channels

Abstract: A hollow, cooled airfoil (14) has a pair of nested, coolant channels (60, 66, 64, 74, 80, 76) therein which carry separate coolant flows back and forth across the span of the airfoil (14) in adjacent parallel paths The coolant in both channels flows from a rearward to forward location within the airfoil (14) allowing the coolant to be ejected from the airfoil near the leading edge (26) through film coolant holes (58)

Experimental Investigation of the Three-Dimensional Flow in an Annular Compressor Cascade

Abstract: A detailed experimental investigation was carried out to examine the influence of blade loading on the three-dimensional flow in an annular compressor cascade. Data were acquired over a range of incidence angles. Included are airfoil and endwall flow visualization, measurement of the static pressure distribution on the flow passage surfaces, and radial-circumferential traverse measurements. The data indicate the formation of a strong vortex near the rear of the blade passage. This vortex transports low-momentum fluid close to the hub toward the blade suction side and seems to be partly responsible for the occurrence of a hub corner stall. The effect of increased loading on the growth of the hub corner stall and its impact on the passage blockage are discussed. Detailed mapping of the blade boundary layer was done to determine the loci of boundary layer transition and flow separation. The data have been compared with results from an integral boundary layer method.

Prediction of Compressor Cascade Performance Using a Navier–Stokes Technique

Abstract: An explicit, time marching, multiple-grid Navier–Stokes technique is demonstrated for the prediction of quasi-three-dimensional turbomachinery compressor cascade performance over the entire incidence range. A numerical investigation has been performed in which the present Navier–Stokes procedure was used to analyze a series of compressor cascade viscous flows for which corresponding experimental data are available. Results from these calculations show that the current viscous flow procedure is capable of predicting cascade profile loss and airfoil pressure distributions with high accuracy. The results from this numerical investigation in the form of comparisons between the predicted profile loss, exit gas angle, and pressure distributions with experimental data are presented in this paper. Results from a grid refinement study are also shown to demonstrate that the Navier–Stokes solutions are grid independent.

Computation of unsteady transonic flows by the solution of Euler equations

Abstract: On utilise un schema de volumes finis pour la discretisation spatiale de la forme integrale des equations d'Euler pour un domaine mobile

Gas turbine vane

Abstract: A gas turbine vane has a vane airfoil defining a cavity extending along the longitudinal direction of the vane airfoil. A guide cylinder is disposed in the cavity to guide coolant fluid supplied from an external source thereof. A plurality of flowing holes are concentrated substantially centrally with respect to the vane airfoil in the longitudinal direction.

Tripple pass cooled airfoil

Abstract: A hollow, cooled airfoil has a pair of nested, coolant channels therein which carry separate coolant flows back and forth across the span of the airfoil in adjacent parallel paths. The coolant in both channels flows from a rearward to forward location within the airfoil allowing the coolant to be ejected from the airfoil near the leading edge through film coolant holes.

Combined translation/pitch motion: a new airfoil dynamic stall simulation

Abstract: The effects of simultaneous velocity and incidence fluctuations on the 2-D aerodynamic behaviour of a NACA 0012 airfoil are investigated in this paper. A new mechanical system allows driving the airfoil in pitching and in fore and aft motions, as well as in a simultaneous combination of these two basic unsteady motions. In response to the simultaneous velocity and incidence variations, the time-dependent lift and drag fluctuations are measured for increasing values of the reduced frequency and amplitude parameters, including dynamic stall conditions. Complementary information on the dynamic stall occurring in combined motion is provided by skin friction and pressure measurements along the airfoil surface.

Airfoil pressure measurements during a blade vortex interaction and a comparison with theory

Abstract: A two-dimensional, airfoil-vortex interaction experiment was conducted to obtain the aerodynamic behavior of the airfoil during a parallel interaction. The vortex was created by pitching a second airfoil placed upstream and parallel to an instrumented test airfoil. Hot-wire anemometer measurements were taken to determine the vortex velocity distribution. The interaction tests were conducted for a counterclockwise vortex passing above a symmetrical airfoil at zero angle of attack

An experimental investigation of flowfield about a multielement airfoil

Abstract: Detailed measurements of mean-flow and turbulence quantities around a multielement airfoil model have been made using pressure and hot-wire probes. The results obtained in two test cases at the chord Reynolds number of 3 million and the freestream Mach number of 0.2 show a number of features of the complex flows that are important in accurate modeling of these flows by numerical methods. Many parts of the shear flow vastly deviate from classical flows, and the interaction with the external potential flow is very strong.

Theoretical prediction of counter-rotating propeller noise

Abstract: A theoretical prediction scheme has been developed for the tone noise generated by a counter-rotation propeller. We start by deriving formulae for the harmonic components of the far acoustic field generated by the thickness and steady loading noise sources. Excellent agreement is shown between theory and measurements. Asymptotic approximation techniques are described which enable us to simplify considerably the complex radiation formulae, whilst retaining all of their important characteristics, and thus save, typically, 95% of computer processing time. Next we derive formulae for the radiated sound field generated by aerodynamic interactions between the blade rows. Here, however, the inputs to the formulae include a knowledge of the fluctuating blade pressure fields which cannot generally be assumed given and must therefore be calculated within the prediction scheme. In the case of viscous wake interactions we consider various models for the wake profile which is written as a series of harmonic gusts. The fluctuating pressure distribution on the downstream blades can then be calculated in the high frequency limit. Comparisons are made between measurements and predictions for a counter-rotation propeller and for rotor/stator interaction on a model fan rig. For potential field interactions we describe the flow fields due to blade circulation and blade thickness in terms of harmonic gusts with the flow assumed incompressible. The blade response is calculated for both finite and semi-infinite airfoils. Some important differences between these two cases are noted in both high and low frequency limits. Predicted noise levels are much improved over those obtained using only the viscous wake model. The inclusion of compressibility, in both flow field and airfoil response calculations, provides a further improvement in the predicted noise levels. The discrepancy between measurements and predictions at this stage is, typically, 2 or 3 dB.

Wingtip vortex turbine

Abstract: A means for extracting rotational energy from the vortex created at aircraft wing tips which consists of a turbine with blades located in the crossflow of the vortex and attached downstream of the wingtip. The turbine 30 has blades 40, 41, 42 and 43 attached to a core 45. When the aircraft is in motion, rotation of core 45 transmits energy to a centrally attached shaft 50. The rotational energy thus generated may be put to use within the airfoil 20 or aircraft fuselage 10.

Passive venting system for modifying cavity flowfields at supersonic speeds

Abstract: The drag of airfoils in transonic flow can be reduced through the use of a passive venting system that employs a porous plate for part of the airfoil upper surface with a vent chamber underneath the porous plate Attention is given to the results obtained with a wind tunnel model employing such a porous floor system. This passive venting system has been used to extend the length/height value before the onset of high drag-producing closed cavity flow at supersonic speeds.

Aerodynamic interactions between a rotor and wing in hover

Abstract: An experimental investigation of rotor/wing aerodynamic interactions in hover is described. The investigation consisted of both a large-scale and a small-scale test. A 0.658-scale V-22 rotor and wing was used in the large-scale test. Wing download, wing surface pressure, rotor performance, and rotor downwash data from the large-scale test are presented. A small-scale experiment was conducted to determine how changes in the rotor/wing geometry affected the aerodynamic interactions. These geometry variations included the distance between the rotor and wing, wing incidence angle, wing flap angle, rotor rotation direction, and configurations both with the rotor axis at the tip of the wing (tilt rotor configuration) and with the rotor axis at the center of the wing (compound helicopter configuration).

Fluid dynamics of airfoils with moving surface boundary-layer control

Abstract: Effectiveness of the moving surface boundary-layer control is assessed with reference to a symmetrical Joukowsky airfoil modified with a leading-edge rotating cylinder. Results of the test program and the numerical models suggest the following: 1) The surface singularity method is essential in modeling the complicated flow. With the inclusion of the boundary-layer correction scheme, it becomes an effective tool for obtaining useful information concerning moving surface boundary-layer control. The predicted pressure distributions are in good agreement with experiment almost up to the point of complete separation from the airfoil surface, except in the separation region, where the prediction of separated boundary layers with flow reversal would require the solution of the full Navier-Stokes equations. 2) The concept of moving surface boundary-layer control appears quite promising. The tests showed a significant improvement in maximum lift and stall characteristics. With cylinder rotation, the flow never separated completely from the upper surface for angles of attack as high as 48 deg. The higher rates of rotation (UC/U>1, Uc = cylinder surface velocity, U = freestream velocity) promoted reattachment of the partially separated flow, giving an increase in lift coefficient by as much as 150% for Ue/U = 4.

Calculation of Flow Over Iced Airfoils

Abstract: Progress toward the development of a method for predicting the flowfield of an iced airfoil is described and shown to offer the prospect of a priori calculations of the effects of ice accretion and roughness on airfoil performance. The approach is based on interaction of inviscid flow solutions obtained by a panel method and improved upon by a finite-difference boundary-layer method which, operating in an inverse mode, incorporates viscous effects including those associated with separated flows. Results are presented for smooth, rough and iced airfoils as a function of angle of attack. Those for smooth and rough airfoils confirm the accuracy of the method and its applicability to surfaces with roughness similar to that associated with insect deposition and some forms of ice. Two procedures have been developed to deal with large ice accretion and their performance is examined and shown to be appropriate to the engineering requirements.

Some numerical and physical aspects of unsteady Navier-Stokes computations over airfoils using dynamic meshes

Abstract: An upwind-biased implicit approximate factorization algorithm is applied to several unsteady flows on dynamic meshes. The thin-layer form of the compressible Navier-Stokes equations is used to solve both laminar and turbulent flows over airfoils pitching about the quarter chord. Numerical aspects of the solutions are investigated, including grid and time step effects. Two methods for determining fluxes - flux-vector splitting and flux-difference splitting - are compared. Flux-difference splitting predicts results more accurately than flux-vector splitting on a coarse mesh, but both methods agree on a fine mesh. Physical aspects of the computations are also examined. An equilibrium turbulent boundary layer model computes generally better unsteady results in comparison with experiment than a nonequilibrium model for the transonic case analyzed. Also, the size and location of the primary shed vortex for an airfoil pitching up at a constant rate is calculated in good agreement with experiment for two pitch rates.

Unsteady vortex-dominated flows around maneuvering wings over a wide range of Mach numbers

Abstract: The problem of unsteady flow around maneuvering wings is solved using the unsteady Euler equations. The unsteady conservative Euler equations are derived for the flow relative motion with respect to a moving (translating and rotating) frame of reference. The resulting equations can handle the most general case for unsteady three-dimensional flow around maneuvering wings or wing-body configurations undergoing six degrees of freedom motion; three translations and three rotations. The equations are solved using two computational schemes; an explicit multistage finite-volume scheme and an implicit approximately factored finite-volume scheme. The computational applications cover two cases. The first case is for a locally conical supersonic flow of rolling oscillation of a sharp-edged delta wing at zero angle of attack. The second case is for a pitching oscillation around a mean angle of attack of a NACA 0012 airfoil in transonic flow.