Search NASA⌕ Search

SEARCH · Search NASA

Results for “AIRFOILS - CASCADES”

Search indexed NASA NTRS and DOE OSTI research on propulsion, heat transfer, battery materials and energy systems. Follow report and document links to the original sources.

Quote a phrase for an exact phrase match. Source license links do not imply unrestricted reuse.

At least 973 records · Page 54
All sourcesNASA NTRSDOE OSTI
NASA NTRS1980

Calculation of the transient motion of elastic airfoils forced by control surface motion and gusts

The time-domain equations of motion of elastic airfoil sections forced by control surface motions and gusts were developed for the case of incompressible flow. Extensive use was made of special functions related to the inverse transform of Theodorsen's function. Approximations for the special cases of zero stream velocity, small time, large and time are given. A numerical solution technique for the solution of the general case is given. Examples of the exact transient response of an airfoil are presented.

Balakrishnan, A. V.↗
NASA NTRS1980

Visualization of the laminar-turbulent transition in the flow over an airfoil using the smoke-wire technique

A flow visualization technique, referred to as the smoke-wire, was used for visualization of the transition of the free shear layer associated with the laminar separation bubble of a NACA 66 sub 3-018 airfoil section at low Reynolds number (50,000-130,000). The smoke-wire technique allows the introduction of fine smoke streaklines into the flow field through the electrical resistive heating of a very fine wire which has been coated with oil and which is located upstream from the leading edge of the airfoil section. Streakline data were collected using both high speed still and motion picture photography.

Batill, S. M.↗
NASA NTRS1980

Aerodynamic coefficients in generalized unsteady thin airfoil theory

Two cases are considered: (1) rigid body motion of an airfoil-flap combination consisting of vertical translation of given amplitude, rotation of given amplitude about a specified axis, and rotation of given amplitude of the control surface alone about its hinge; the upwash for this problem is defined mathematically; and (2) sinusoidal gust of given amplitude and wave number, for which the upwash is defined mathematically. Simple universal formulas are presented for the most important aerodynamic coefficients in unsteady thin airfoil theory. The lift and moment induced by a generalized gust are evaluated explicitly in terms of the gust wavelength. Similarly, in the control surface problem, the lift, moment, and hinge moments are given as explicit algebraic functions of hinge location. These results can be used together with any of the standard numerical inversion routines for the elementary loads (pitch and heave).

Williams, M. H.↗
NASA NTRS1980

Shockless airfoils for wings, compressors, and turbines

The major aspects of a code for the design of shockless airfoils in two-dimensional transonic flow are reviewed. The code is based on a method of analytic continuation into the domain of two complex characteristic variables, which makes it possible to associate a shockless airfoil with a given pressure distribution. This procedure is compared with another more direct method based on a two-dimensional analysis code that has had wide acceptance by the aircraft industry.

Garabedian, P. R.↗
NASA NTRS1980

Illustration of airfoil shape effect on forward-swept wing divergence

A static aeroelastic analysis is presented of the divergence of untapered wings with conventional and supercritical airfoil sections at sweep angles of zero and -15 deg. One bending and one torsion mode were employed for a uniform rectangular cantilevered beam with the elastic axis at midchord, and calculations were based on a two-dimensional differential equations formulation in the structural coordinate system and in simple strip theory. A minimum divergence speed in the transonic range is obtained which is associated with the rearward shift of the aerodynamic center, and a 17% difference in minimum divergence dynamic pressure is found between a supercritical and a conventional wing. It is noted that although the strip method employed allows the assessment of the sensitivity of airfoil shapes to divergence, three-dimensional transonic aerodynamic methods should be used to predict wing divergence characteristics.

Bland, S. R.↗
NASA NTRS1982

Algorithm for Unsteady Potential Flow About Airfoils

Implicit finite-difference scheme efficiently computes unsteady potential flow about airfoils. Formulation uses density and velocity potential as dependent variables. Conservation form is retained to assure that shock wave location and speed are computed correctly. Scheme fills need for method to calculate efficiently unsteady potential flow about airfoils and to predict flutter and other unsteady aeroelastic phenomena in transonic flow regimes.

Chipman, R.↗
NASA NTRS1980

A computer program for the design and analysis of low-speed airfoils, supplement

Three new options were incorporated into an existing computer program for the design and analysis of low speed airfoils. These options permit the analysis of airfoils having variable chord (variable geometry), a boundary layer displacement iteration, and the analysis of the effect of single roughness elements. All three options are described in detail and are included in the FORTRAN IV computer program.

Eppler, R.↗
NASA NTRS1981

Analysis of turbulent flow about an isolated airfoil using a time dependent Navier-Stokes procedure

The procedure solves the Navier-Stokes equations by the consistently split linearized block implicit method of Briley and McDonald in a body fitted coordinate system. The procedure is described and results are presented for flow about an airfoil whose incidence changes from 6 degrees to 19 degrees at a Reynolds number of one million and Mach number of 0.2. In addition, the unsteady flow about an airfoil held at a constant 19 degree incidence is examined and compared to data.

Shamroth, S. J.↗
NASA NTRS1981

TAIR: A transonic airfoil analysis computer code

The operation of the TAIR (Transonic AIRfoil) computer code, which uses a fast, fully implicit algorithm to solve the conservative full-potential equation for transonic flow fields about arbitrary airfoils, is described on two levels of sophistication: simplified operation and detailed operation. The program organization and theory are elaborated to simplify modification of TAIR for new applications. Examples with input and output are given for a wide range of cases, including incompressible, subcritical compressible, and transonic calculations.

Dougherty, F. C.↗
NASA NTRS1981

Prediction of boundary-layer characteristics of an oscillating airfoil

The evolution of unsteady boundary layers on oscillating airfoils is investigated by solving the governing equations by the Characteristic Box scheme. The difficulties associated with computing the first profile on a given time line, and the velocity profiles with partial flow reversal are solved. A sample calculation is performed for an external velocity distribution typical of those found near the leading edge of thin airfoils. The viability of the calculation procedure is demonstrated.

Cebeci, T.↗
NASA NTRS1981

Improving turbine engine compressor performance retention through airfoil coatings

In order to evaluate the potential effectiveness of coatings in limiting erosive damage to compressor airfoils, an effort was initiated to evaluate candidate coatings for substrate alloys typically used in commercial engine high compressor blades. Laboratory and rig erosion testing of plasma deposited and diffusion coatings described in this paper have shown the potential of a two to four fold improvement in erosion life. The selective application of these coatings to approximately the outer third of the airfoil avoids coating the fatigue critical region of the blade, thus providing erosion resistance potentially without compromising the fatigue strength of the blade. Both the plasma and the diffusion coatings also offer the advantage of low initial cost and a multi-source production base.

Friedrich, L. A.↗
NASA NTRS1981

Icing tunnel tests of a glycol-exuding porous leading edge ice protection system on a general aviation airfoil

Tests were conducted in the Icing Research Tunnel at the NASA Lewis Research Center to determine the characteristics of an ice protection system that distributes a glycol solution onto the leading edge of an airfoil through a porous surface material. Minimum fluid flow rates required to achieve anti-icing (no ice formation) were determined for various flight conditions and angles of attack. The ability of the system to remove ice formed on the airfoil before system activation was also investigated.

Kohlman, D. L.↗
NASA NTRS1980

Design considerations of advanced supercritical low drag suction airfoils

Supercritical low drag suction laminar flow airfoils were laid out for shock-free flow at design freestream Mach = 0.76, design lift coefficient = 0.58, and t/c = 0.13. The design goals were the minimization of suction laminarization problems and the assurance of shock-free flow at freestream Mach not greater than design freestream Mach (for design lift coefficient) as well as at lift coefficient not greater than design lift coefficient (for design freestream Mach); this involved limiting the height-to-length ratio of the supersonic zone at design to 0.35. High design freestream Mach numbers result with extensive supersonic flow (over 80% of the chord) on the upper surface, with a steep Stratford-type rear pressure rise with suction, as well as by carrying lift essentially in front- and rear-loaded regions of the airfoil with high static pressures on the carved out front and rear lower surface.

Pfenninger, W.↗
NASA NTRS1981

Numerical simulation of steady transonic flow about airfoils

A computer code has been developed that couples a fast transonic full-potential AF2 solver with both an efficient integral boundary-layer method and a viscous wedge approximation of the shock/boundary-layer interaction. The efficiency of the coupled analysis methods and the method of coupling has resulted in a uniquely efficient analysis tool. The airfoil geometry is modified by the displacement thickness before the shock and the displacement thickness plus the viscous wedge thickness after the shock by considering the viscous effects as an equivalent transpiration boundary condition. The flow about conventional and supercritical airfoils under moderately strong shock situations has been calculated. Comparisons with experimental data indicate that this viscous correction method has improved the accuracy of the full-potential analysis. Furthermore, the computer time required to obtain a converged solution has been reduced.

Lee, S. C.↗
NASA NTRS1981

On the structure of turbulent wakes and merging shear layers of multielement airfoils

The structure of the attached turbulent flow in the vicinity of a NACA 4412 airfoil equipped with a single-slotted flap was studied. The airfoil/flap configuration was tested at a Mach number of 0.06 and a Reynolds number of 1.3 x 10 to the 6th in a 7- by 10-Foot Wind Tunnel. Surface-pressure measurements were made on the main airfoil and on the flap. Detailed measurements, obtained using a high-spatial-resolution laser Doppler anemometer, were made of the mean velocity flow field and of the second-order statistical quantities (Reynolds stresses) in the boundary layers, wakes, and merging shear layers. The experimental observations are compared with theoretical predictions of pressure, mean velocity, and Reynolds stress.

Olson, L. E.↗
NASA NTRS1981

Application of unsteady airfoil theory to rotary wings

A clarification is presented on recent work concerning the application of unsteady airfoil theory to rotary wings. The application of this theory may be seen as consisting of four steps: (1) the selection of an appropriate unsteady airfoil theory; (2) the resolution of that velocity which is the resultant of aerodynamic and dynamic velocities at a point on the elastic axis into radial, tangential and perpendicular components, and the angular velocity of a blade section about the deformed axis; (3) the expression of lift and pitching moments in terms of the three components; and (4) the derivation of explicit expressions for the components in terms of flight velocity, induced flow, rotor rotational speed, blade motion variables, etc.

Kaza, K. R. V.↗
NASA NTRS1953

Analytical Investigation of Icing Limit for Diamond-Shaped Airfoil in Transonic and Supersonic Flow

Calculations have been made for the icing limit of a diamond airfoil at zero angle of attack in terms of the stream Mach number, stream temperature, and pressure altitude. The icing limit is defined as a wetted-surface temperature of 320 F and is related to the stream conditions by the method of Hardy. The results show that the point most likely to ice on the airfoil lies immediately behind the shoulder and is subject to possible icing at Mach numbers as high as 1.4.

Callaghan, Edmund E.↗
NASA NTRS1952

Preliminary Results of Cyclical De-Icing of a Gas-Heated Airfoil

An NACA 65(sub 1)-212 airfoil of 8-foot chord was provided with a gas-heated leading edge for investigations of cyclical de-icing. De-icing was accomplished with intermittent heating of airfoil segments that supplied hot gas to chordwise passages in a double-skin construction. Ice removal was facilitated by a spanwise leading-edge parting strip which was continuously heated from the gas-supply duct. Preliminary results demonstrate that satisfactory cyclical ice removal occurs with ratios of cycle time to heat-on period (cycle ratio) from 10 to 26. For minimum runback, efficient ice removal, and minimum total heat input, short heat-on periods of about 15 seconds with heat-off periods of 260 seconds gave the best results. In the range of conditions investigated, the prime variables in the determination of the required heat input for cyclical ice removal were the air temperature and the cycle ratio; heat-off period, liquid water content, airspeed, and angle of attack had only secondary effects on heat input rate.

Gray, V. H.↗