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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.

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At least 343 records · Page 19
All sourcesNASA NTRSDOE OSTI
NASA NTRS2016

Ice Accretions and Full-Scale Iced Aerodynamic Performance Data for a Two-Dimensional NACA 23012 Airfoil

This report documents the data collected during the large wind tunnel campaigns conducted as part of the SUNSET project (StUdies oN Scaling EffecTs due to ice) also known as the Ice-Accretion Aerodynamics Simulation study: a joint effort by NASA, the Office National d'Etudes et Recherches Aérospatiales (ONERA), and the University of Illinois. These data form a benchmark database of full-scale ice accretions and corresponding ice-contaminated aerodynamic performance data for a two-dimensional (2D) NACA 23012 airfoil. The wider research effort also included an analysis of ice-contaminated aerodynamics that categorized ice accretions by aerodynamic effects and an investigation of subscale, low- Reynolds-number ice-contaminated aerodynamics for the NACA 23012 airfoil. The low-Reynolds-number investigation included an analysis of the geometric fidelity needed to reliably assess aerodynamic effects of airfoil icing using artificial ice shapes. Included herein are records of the ice accreted during campaigns in NASA Glenn Research Center's Icing Research Tunnel (IRT). Two different 2D NACA 23012 airfoil models were used during these campaigns; an 18-in. (45.7-cm) chord (subscale) model and a 72-in. (182.9-cm) chord (full-scale) model. The aircraft icing conditions used during these campaigns were selected from the Federal Aviation Administration's (FAA's) Code of Federal Regulations (CFR) Part 25 Appendix C icing envelopes. The records include the test conditions, photographs of the ice accreted, tracings of the ice, and ice depth measurements. Model coordinates and pressure tap locations are also presented. Also included herein are the data recorded during a wind tunnel campaign conducted in the F1 Subsonic Pressurized Wind Tunnel of ONERA. The F1 tunnel is a pressured, high- Reynolds-number facility that could accommodate the full-scale (72-in. (182.9-cm) chord) 2D NACA 23012 model. Molds were made of the ice accreted during selected test runs of the full-scale model in the IRT. From these molds, castings were made that closely replicated the features of the accreted ice. The castings were then mounted on the full-scale model in the F1 tunnel, and aerodynamic performance measurements were made using model surface pressure taps, the facility force balance system, and a large wake rake designed specifically for these tests. Tests were run over a range of Reynolds and Mach numbers. For each run, the model was rotated over a range of angles-of-attack that included airfoil stall. The benchmark data collected during these campaigns were, and continue to be, used for various purposes. The full-scale data form a unique, ice-accretion and associated aerodynamic performance dataset that can be used as a reference when addressing concerns regarding the use of subscale ice-accretion data to assess full-scale icing effects. Further, the data may be used in the development or enhancement of both ice-accretion prediction codes and computational fluid dynamic codes when applied to study the effects of icing. Finally, as was done in the wider study, the data may be used to help determine the level of geometric fidelity needed for artificial ice used to assess aerodynamic degradation due to aircraft icing. The structured, multifaceted approach used in this research effort provides a unique perspective on the aerodynamic effects of aircraft icing. The data presented in this report are available in electronic form upon formal approval by proper NASA and ONERA authorities.

Addy, Harold E., Jr.↗
NASA NTRS2018

Experimental Study of Noise Shielding by a NACA 0012 Airfoil

The effects of sound source location, Mach number and angle of attack on the shielding of a laser-induced sound source by a NACA 0012 airfoil are examined. The sound source is a small plasma generated by a high energy, laser beam focused to a point. In-flow microphone measurements are acquired in the midspan plane of the airfoil over a broad range of streamwise stations, and shielding levels are calculated over different frequency ranges from the measurements acquired with and without the airfoil installed. Shielding levels are shown to increase as the source is positioned closer to the mid-chord of the airfoil, and to significantly decrease with increasing flow Mach number, except when the source is positioned near the leading edge of the airfoil. Both with and without flow, changes in angle of attack are associated with a corresponding shift of the shadow region. Finally, the effects of multipath signals, observer distance and signal scatter on the measured shielding levels are discussed.

Hutcheson, Florence V.↗
NASA NTRS2019

Airfoil Selection for Mars Rotor Applications

This short white paper is aimed at providing an overview of design considerations for airfoil choices for rotor applications in the Martian atmosphere, at very low chord-based Reynolds number flows, around Rec = O (103 -104). The low Reynolds number typical of rotorcraft operation in the Martian atmosphere reduces the rotor lifting force and efficiency, which is only partially compensated for by a lower gravity on Mars. Additionally, the low temperature and largely CO2 based atmosphere of Mars compound the overall aerodynamic problem by resulting in a lower speed of sound, further constraining rotor operation in the Martian atmosphere by limiting the maximum rotor tip speed possible so as not to exceed an acceptable tip Mach number. In light of the expected reduced rotor efficiency, evaluating airfoils for compressible, low-Reynolds number Mars rotor applications is key. Prior research on airfoil optimization and performance evaluation at low Reynolds numbers, especially in the compressible regime, is scarce and deserves further understanding. Specifically, the proposed goal stemming from this overview is to develop airfoils tailored to the unique demands of the second generation of Mars rotorcraft, i.e. the Mars Science Helicopter (MSH). This research focuses on the airfoil performance at low Reynolds numbers and hopes to add to the work performed by, amongst others, Kroo et al., Kunz and Kroo, Oyama and Fujii, Anyoji et al.

Koning, Witold J. F.↗
NASA NTRS2020

Robust Conceptual Design of Transonic Airfoils

This paper describes an integrated, multi-fidelity analysis and heuristic design approach that can be used to derive initial airfoil designs for transonic flight. If successful, the final result is a geometry that can be expected to produce reasonable aerodynamic performance when used with higher order analysis methods. A key aspect of the methodology is the use of a “sonic-plateau” pressure distribution as the target distribution for inverse design. The sonic- plateau distribution is easily parameterized and has the advantage of automatically resulting in a smooth airfoil shape without any discontinuities built into the surface due to the presence of a shock in the target pressure distribution. Inverse design is performed on each airfoil using a parametrically defined pressure distribution at a reduced lift coefficient and Mach number from the design condition. The methodology is demonstrated by designing an airfoil at 38% of the wing semispan for a 737-200-like aircraft. The demonstration problem shows that the methodology is able to achieve rapid and robust convergence to the solution. The calculated designed airfoil was found to be sufficiently higher than Mach number, and the maximum thickness was close to the targeted value.

Olson, Erik D.↗
NASA NTRS1927

Experiments on airfoils with trailing edge cut away

Airfoils with their trailing edge cut away are often found on aircraft, as the fins on the hulls of flying boats and the central section of the wings for affording better visibility. It was therefore of some interest to discover the effect of such cutaways on the lift and drag and on the position of the center of pressure. For this purpose, systematic experiments were performed on two different airfoils, a symmetrical airfoil and an airfoil of medium thickness, with successive shortenings of their chords.

AIRFOILS↗
NASA NTRS1941

Wind-Tunnel Investigation of Perforated Split Flaps for Use as Dive Brakes on a Tapered NACA 23012 Airfoil

Aerodynamic characteristics of a tapered NACA 23012 airfoil with single and double perforated split flaps have been determined in the NACA 7- by 10-foot wind tunnel. Dynamic pressure surveys were made behind the airfoil at the approximate location of the tail in order to determine the extent and location of the wake for several of the flap arrangements. In addition, computations have been made of an application of perforated double split flaps for use as fighter brakes. The results indicated that single or double perforated split flaps may be used to obtain satisfactory dive control without undue buffeting effects and that single or double perforated split flaps may also be used as fighter brakes. The perforated split flaps had approximately the same effects on the aerodynamic and wake characteristics of the tapered airfoil as on a comparable rectangular airfoil.

WING FLAPS, SPLIT - PERFORATED↗
NASA NTRS1943

Wind-Tunnel Investigation of a Low-Drag Airfoil Section with a Double Slotted Flap

Tests were made of an 0.309-chord double-slotted flap on an NACA 65, 3-118, a equals 1.0 airfoil section to determine drag, lift, and pitching-moment characteristics for a range of flap deflections. Results indicate that combination of a low-drag airfoil and a double-slotted flap, of which the two parts moved as a single unit, gave higher maximum lift coefficients than have been obtained with plain, split, or slotted flaps on low-drag airfoils. Pitching moments were comparable to those obtained with other high-lift devices on conventional airfoils for similar lift coefficients.

AIRFOILS - NACA 65,3-118-CHAR , MOMENTS & ORDINATE↗
NASA NTRS1947

Experimental and calculated characteristics of three wings of NACA 64-210 and 65-210 airfoil sections with and without 2 degree washout

Report presents the results of an investigation conducted to determine some of the effects of airfoil section and washout on the experimental and calculated characteristics of 10-percent-thick wings. Three wings of aspect ratio 9 and ratio of root chord to tip chord 2.5 were tested. One wing had NACA 64-210 sections and 2 degree washout, the second had NACA 65-210 sections and 2 degree washout, and the third had NACA 65-210 sections and 0 degree washout. It was found that the experimental characteristics of the wings could be satisfactorily predicted from calculations based upon two-dimensional data when the airfoil contours of the wings conformed to the true airfoil sections with the same high degree of accuracy as the two-dimensional models.

AIRFOILS - NACA 65-210 - CHAR & MOMENTS LOW-DRAG↗
NASA NTRS1970

Leading-edge pressure measurements of airfoil vortex interaction

Experimental pressure-differential measurements made at 10% chord of an airfoil-vortex interaction are presented. A line vortex was oscillated over an airfoil perpendicular to the span and parallel to the chord. The pressure time history was recorded in order to show the sharp pressure pulses resulting from the bursting of the vortex core as it impinges upon the airfoil. Results for various vortex sizes and free stream velocities were obtained. Measurements were also made when the airfoil was yawed to the line vortex. Maximum pressure differences were observed to occur in phase across the blade even with yaw, and were directly proportional to the square of the free stream velocity. The maximum dynamic pressure coefficients obtained were as high as 1.0 when vortex bursting occurred.

Walsh, R. G., Jr.↗
NASA NTRS1972

A finite-difference method for transonic airfoil design.

This paper describes an inverse method for designing transonic airfoil sections or for modifying existing profiles. Mixed finite-difference procedures are applied to the equations of transonic small disturbance theory to determine the airfoil shape corresponding to a given surface pressure distribution. The equations are solved for the velocity components in the physical domain and flows with embedded shock waves can be calculated. To facilitate airfoil design, the method allows alternating between inverse and direct calculations to obtain a profile shape that satisfies given geometric constraints. Examples are shown of the application of the technique to improve the performance of several lifting airfoil sections. The extension of the method to three dimensions for designing supercritical wings is also indicated.

Steger, J. L.↗
NASA NTRS1973

Steady inviscid transonic flows over planar airfoils: A search for a simplified procedure

A finite difference procedure based upon a system of unsteady equations in proper conservation form with either exact or small disturbance steady terms is used to calculate the steady flows over several classes of airfoils. The airfoil condition is fulfilled on a slab whose upstream extremity is a semi-circle overlaying the airfoil leading edge circle. The limitations of the small disturbance equations are demonstrated in an extreme example of a blunt-nosed, aft-cambered airfoil. The necessity of using the equations in proper conservation form to capture the shock properly is stressed. Ability of the steady relaxation procedures to capture the shock is briefly examined.

Magnus, R.↗
NASA NTRS1973

An analytical study for the design of advanced rotor airfoils

A theoretical study has been conducted to design and evaluate two airfoils for helicopter rotors. The best basic shape, designed with a transonic hodograph design method, was modified to meet subsonic criteria. One airfoil had an additional constraint for low pitching-moment at the transonic design point. Airfoil characteristics were predicted. Results of a comparative analysis of helicopter performance indicate that the new airfoils will produce reduced rotor power requirements compared to the NACA 0012. The hodograph design method, written in CDC Algol, is listed and described.

Kemp, L. D.↗
NASA NTRS1973

The effects of leading-edge serrations on reducing flow unsteadiness about airfoils.

High frequency surface pressure measurements were obtained from wind-tunnel tests over the Reynolds number range 1.2 x 1,000,000 to 6.2 x 1,000,000 on a rectangular wing of NACA 63-009 airfoil section. A wide selection of leading-edge serrations were also added to the basic airfoil. Under a two-dimensional laminar bubble very close to the leading edge of the basic airfoil there is a large peak in rms pressure, which is interpreted as an oscillation in size and position of the bubble. The serrations divide the bubble into segments and reduce the peak rms pressures. A low Reynolds number flow visualization test on a hydrofoil in water was also conducted. A von Karman vortex street was found trailing from the rear of the foil. Its frequency is at a much lower Strouhal number than in the high Reynolds number experiment, and is related mathematically to the airfoil trailing-edge and boundary-layer thicknesses.

Schwind, R. G.↗
NASA NTRS1974

Unsteady lift forces on highly cambered airfoils moving through a gust

An unsteady airfoil theory in which the flow is linearized about the steady potential flow of the airfoil is presented. The theory is applied to an airfoil entering a gust. After transformation to the W-plane, the problem is formulated in terms of a Poisson's equation. The solutions are expanded in a Fourier-Bessel series. The theory is applied to a circular arc with arbitrary camber. Closed form expressions for the velocity and pressure on the surface of the airfoil are obtained. The unsteady aerodynamic forces are then calculated and shown to contain two terms. One in an explicit closed analytical form represents the contribution of the oncoming vortical disturbance, the other depends on a single quadrature and accounts for the effect of the wake.

Atassi, H.↗
NASA NTRS1975

Transonic airfoil flowfield analysis using Cartesian coordinates

A numerical technique for analyzing transonic airfoils is presented. The method employs the basic features of Jameson's iterative solution for the full potential equation, except that Cartesian coordinates are used rather than a grid which fits the airfoil, such as the conformal circle-plane or 'sheared parabolic' coordinates which were used previously. Comparison with previous results shows that it is not necessary to match the computational grid to the airfoil surface, and that accurate results can be obtained with a Cartesian grid for lifting supercritical airfoils.

Carlson, L. A.↗
NASA NTRS1975

Surface pressure and wake flow fluctuations in a supercritical airfoil flowfield

Nonsteady features of a Whitcomb-type supercritical airfoil flowfield were studied in a series of transonic wind tunnel experiments. Data consisted of mean and fluctuating pressures on the airfoil and in the free stream, velocity fluctuations in the wake region, and airfoil vibrations. Fluctuation data were analyzed statistically for intensity, frequency content, and spatial coherence; variations in these parameters were correlated with changes in the mean airfoil flowfield. Surface pressure fluctuation intensity was influenced primarily by the location and motion of the upper-surface shock, the existence of separation, and (downstream of the shock) the development and upstream propagation of 'shocklets'. Power spectra of pressure fluctuations showed characteristics differences up- and downstream of the shock. Chordwise and spanwise coherences were considerably reduced in the presence of separation. Downwash fluctuations at the edge of the wake were noticeably affected only by the development of trailing-edge flow separation.

Roos, F. W.↗
NASA NTRS1975

Starting vortex, separation bubbles and stall - A numerical study of laminar unsteady flow around an airfoil

The stalling characteristics of an airfoil in a laminar viscous incompressible fluid are investigated. The governing equations in terms of the vorticity and stream function are solved using an implicit finite-difference scheme and point successive relaxation procedure. The development of the impulsively started flow, the initial generation of circulation, and the behavior of the forces at large times are studied. Following the impulsive start, the lift is at first very large and then it rapidly drops. The subsequent growth of circulation and lift is associated with the starting vortex. After incipient separation, the lift increases owing to enlargement of the separation bubble and intensification of the flow rotation in it. The extension of this bubble beyond the trailing edge causes it to rupture and brings about the stalling characteristics of the airfoil. Subsequently, new bubbles are formed near the upper surface of the airfoil and are swept away. The behavior of the lift acting on the airfoil is explained in terms of the strength and sense of these bubbles.

Mehta, U. B.↗
NASA NTRS1975

Transonic airfoil analysis and design using Cartesian coordinates

An inverse numerical technique for designing transonic airfoils having a prescribed pressure distribution is presented. The method uses the full potential equation, inverse boundary conditions, and Cartesian coordinates. It includes simultaneous airfoil update and utilizes a direct-inverse approach that permits a logical method for controlling trailing edge closure. The method can also be used for the analysis of flowfields about specified airfoils. Comparison with previous results shows that accurate results can be obtained with a Cartesian grid. Examples show the application of the method to design aft-cambered and other airfoils specifically for transonic flight.

Carlson, L. A.↗