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At least 19 records

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NASA NTRS2016

Active Flow Control Using Sweeping Jet Actuators on a Semi-Span Wing Model

Wind tunnel experiments were performed using active flow control on an unswept semispan wing model with a 30% chord trailing edge flap to aid in the selection of actuators for a planned high Reynolds number experiment. Two sweeping jet actuator sizes were investigated to determine the influence of actuator size on the active flow control system efficiency. Sweeping jet actuators with orifice sizes of 1 mm x 2 mm and 2 mm x 4 mm were selected because of the differences in actuator jet sweep angle. The parameters that were varied include actuator momentum, freestream velocity, and trailing edge flap deflection angle. Steady and unsteady pressure data, Particle Image Velocimetry data, and force and moment data were acquired to assess the performance of the two actuators. In addition to the wind tunnel experiments, benchtop studies of the actuators were performed to characterize the jets produced by each actuator. Benchtop investigations of the smaller actuator reveal that the jet exiting the actuator has a reduced sweep angle compared to published data for larger versions of this type of actuator. The larger actuator produces an oscillating jet that attaches to the external di↵user walls at low supply pressures and produces the expected sweep angles. The AFC results using the smaller actuators show that while the actuators can control flow separation, the selected spacing of 3.3 cm may be too large due to the reduced sweep angle. In comparison, the spacing for the larger actuators, 6.6 cm, appears to be optimal for the Mach numbers investigated. Particle Image Velocimetry results are presented and show how the wall jets produced by the actuators cause the flow to attach to the flap surface.

Melton, LaTunia Pack↗

NASA NTRS1974

Neutral and ion exosphere models for lunar hydrogen and helium

A general neutral exosphere model, which includes density and temperature variations at the exobase, is applied to the moon to obtain surface and radial density distributions for H, H2, and He. It is assumed that the source for these constituents derives from accretion of solar wind ions. The surface distributions are determined by requiring that the sum of the neutral and solar wind ion fluxes for a given constituent vanish at all points on the surface. On this basis, maximum dayside surface densities for H, H2, and He and maximum nightside surface densities for H, H2, and He are obtained that are consistent with either measured values or upper limits. In addition, model ion density distributions for H2(+) and He(+) are constructed. This ion exosphere is produced by ionization of the neutral exosphere in the solar wind, which efficiently sweeps the ions past or onto the lunar surface. Saturated H2(+) and He(+) densities ranging from about .001 to .015 per cu cm and .00003 to .0004 per cu cm over 1.5 to 3 selenocentric radii on the dayside, respectively.

Hartle, R. E.↗

NASA NTRS2018

Active Flow Control for Trailing Edge Flap Separation

Active flow control (AFC) in the form of sweeping jet (SWJ) excitation and discrete steady jet excitation is used to control the flow separation on an NACA 0015 semispan wing with a deflected, simple-hinged, trailing edge flap. This geometry has been the focus of several recent publications that investigated methods to improve the efficiency of sweeping jet actuators. In the current study, the interaction of the AFC excitation with the separated flowfields present at several flap deflection angles was examined. Previous studies with this model have been limited to a maximum flap deflection angle of 40°. The flap deflection range was extended to 60! because systems studies have indicated that a high-lift system with simple-hinged flaps may require larger flap deflections than the Fowler flaps found on most high-lift systems. The results obtained at flap deflection angles of 20°, 40°, and 60° are presented and compared. Force and moment data, Particle Image Velocimetry (PIV) data, and steady and unsteady surface pressure data are used to describe the flowfield with and without AFC. With a flap deflection of 60°, increasing the SWJ actuator momentum at the flap shoulder increased lift due to an increase in circulation but did not completely eliminate the recirculation region above the flap surface. AFC using the discrete steady jet actuators of this study increased lift as well but required more mass flow than the SWJ actuators and had a detrimental effect on lift at the highest mass flow level tested. PIV results showed that the angle between the excitation and the flap surface was not optimal for attaching the separated shear layer.

Melton, LaTunia Pack↗

NASA NTRS1995

Effect of individual blade control on noise radiation

In a joint research program of NASA Ames Research Center, ZF Luftfahrttechnik, the German Aerospace Research Establishment (DLR), and EUROCOPTER Deutschland, a wind tunnel test was performed to evaluate the effects of Individual Blade Control (IBC) on rotor noise. This test was conducted in the 40x80 ft wind tunnel at NASA Ames Research Center, utilizing a full scale MBB-BO 105 four-bladed rotor system. Three microphones were installed for determination of the radiated noise, two of them on a moveable traverse below the advancing blade side and one in a fixed location below the retreating side. Acoustic results are presented for flight conditions with Blade-Vortex-Interaction (BVI) noise radiation. High noise level reductions were measured for single harmonic control inputs. In addition to the single harmonic inputs, multi-harmonic inputs were evaluated by superimposing 2/rev to 6/rev harmonics. For the first time the efficiency of sharp wavelets (60 deg and 90 deg width) on acoustic noise were measured. In order to achieve an adequate wavelet shape at the blade tip, corrections were made to account for the blade torsional behavior. In parallel with the acoustic measurements, vibratory loads were measured during the BVI flight condition to correlate the effects of IBC on noise and vibrations. It is shown how noise levels and vibrations are affected by specific IBC control inputs. In addition, correlations are made between noise levels and acoustic time histories with IBC phase and amplitude variations. For one IBC input mode with high noise reducing efficiency, a sweep of the moveable microphone traverse below the advancing side shows the effect on BVI noise directivity.

Swanson, S. M.↗

NASA NTRS2001

Estimation of Aircraft Unsteady Aerodynamic Parameters from Dynamic Wind Tunnel Testing

Improved aerodynamic mathematical models, for use in aircraft simulation or flight control design, are required when representing nonlinear unsteady aerodynamics. A key limitation of conventional aerodynamic models is the inability to map frequency and amplitude dependent data into the equations of motion directly. In an effort to obtain a more general formulation of the aerodynamic model, researchers have been led to a parallel requirement for more general testing methods. Testing for a more comprehensive model can lead to a very time consuming number of tests especially if traditional single frequency harmonic testing is attempted. This paper presents an alternative to traditional single frequency forced-oscillation testing by utilizing Schroeder sweeps to efficiently obtain the frequency response of the unsteady aerodynamic model. Schroeder inputs provide signals with a flat power spectrum over a specified frequency band. For comparison, experimental results using the traditional single-frequency inputs are also considered. A method for data analysis to determine an adequate unsteady aerodynamic model is presented. Discussion of associated issues that arise during this type of analysis and comparison of results using traditional single frequency analysis are provided.

Murphy, Patrick C.↗

NASA NTRS1985

Marching iterative methods for the parabolized and thin layer Navier-Stokes equations

Downstream marching iterative schemes for the solution of the Parabolized or Thin Layer (PNS or TL) Navier-Stokes equations are described. Modifications of the primitive equation global relaxation sweep procedure result in efficient second-order marching schemes. These schemes take full account of the reduced order of the approximate equations as they behave like the SLOR for a single elliptic equation. The improved smoothing properties permit the introduction of Multi-Grid acceleration. The proposed algorithm is essentially Reynolds number independent and therefore can be applied to the solution of the subsonic Euler equations. The convergence rates are similar to those obtained by the Multi-Grid solution of a single elliptic equation; the storage is also comparable as only the pressure has to be stored on all levels. Extensions to three-dimensional and compressible subsonic flows are discussed. Numerical results are presented.

Israeli, M.↗

NASA NTRS1986

System design and integration of the large-scale advanced prop-fan

In recent years, considerable attention has been directed toward improving aircraft fuel consumption. Studies have shown that blades with thin airfoils and aerodynamic sweep extend the inherent efficiency advantage that turboprop propulsion systems have demonstrated to the higher speed to today's aircraft. Hamilton Standard has designed a 9-foot diameter single-rotation Prop-Fan. It will test the hardware on a static test stand, in low speed and high speed wind tunnels and on a research aircraft. The major objective of this testing is to establish the structural integrity of large scale Prop-Fans of advanced construction, in addition to the evaluation of aerodynamic performance and the aeroacoustic design. The coordination efforts performed to ensure smooth operation and assembly of the Prop-Fan are summarized. A summary of the loads used to size the system components, the methodology used to establish material allowables and a review of the key analytical results are given.

Huth, B. P.↗

NASA NTRS1985

Marching iterative methods for the Parabolized and Thin Layer Navier-Stokes equations

Israeli, M.↗

NASA NTRS1987

Numerical study of laminar and turbulent flows inside a turnaround duct with and without guide vanes

The purpose of this study is to examine in detail incompressible laminar and turbulent flows inside a turnaround duct with and without guide vanes and to investigate the effects of vanes on the flow characteristics. To perform this study, an implicit finite difference code cast in general curvilinear coordinates is further developed. The code is based on the method of pseudo-compressibility and utilize ADI or implicit approximate factorization algorithm to achieve computational efficiency. Method of segmental sweeping is developed to overcome the multiple-zone problem due to imposition of guide vanes. In the present study, several test cases have been computed. These test cases include laminar and turbulent flows inside a turnaround duct without and with two or three guide vanes. The study reveals that: (1) there exists large recirculation zones inside the duct if no vanes are present; (2) properly shaped and positional guide vanes are effective in eliminating flow separation; and (3) laminar and turbulent flows have similar flow features. But turbulent flow has less total pressure drop.

Lin, S.-J.↗

NASA NTRS1990

The aerodynamic design of the oblique flying wing supersonic transport

The aerodynamic design of a supersonic oblique flying wing is strongly influenced by the requirement that passengers must be accommodated inside the wing. It was revealed that thick oblique wings of very high sweep angle can be efficient at supersonic speeds when transonic normal Mach numbers are allowed on the upper surface of the wing. The goals were motivated by the ability to design a maximum thickness, minimum size oblique flying wing. A 2-D Navier-Stokes solver was used to design airfoils up to 16 percent thickness with specified lift, drag and pitching moment. A new method was developed to calculate the required pressure distribution on the wing based on the airfoil loading, normal Mach number distribution and theoretical knowledge of the minimum drag of oblique configurations at supersonic speeds. The wing mean surface for this pressure distribution was calculated using an inverse potential flow solver. The lift to drag ratio of this wing was significantly higher than that of a comparable delta wing for cruise speeds up to Mach 2.

Vandervelden, Alexander J. M.↗

NASA NTRS1985

An Attached Flow Design of a Noninterferring Leading Edge Extension to a Thick Delta Wing

An analytical procedure for the determination of the shape of a Leading-Edge Extension (LEE) which satisfies design criteria, including especially noninterference at the wing design point, has been developed for thick delta wings. The LEE device best satisfying all criteria is designed to be mounted on a wing along a dividing stream surface associated with an attached flow design lift coefficient (C L,d ) of greater than zero. This device is intended to improve the aerodynamic performance of transonic aircraft at C L greater than C L,d system emanating from the LEE leading edge. In order to quantify this process a twisted and cambered thick delta wing was chosen for the initial application of this design procedure. Appropriate computer codes representing potential and vortex flows were employed to determine the dividing stream surface at C L,d and an optimized LEE planform shape at C L greater than C L,d , respectively. To aid in the LEE selection, the aerodynamic effectiveness of 36 planforms was investigated at C L greater than C L,d . This study showed that reducing the span of the candidate LEEs has the most detrimental effect on overall aerodynamic efficiency, regardless of the shape or area. Furthermore, for a fixed area, constant-chord LEE candidates were relatively more efficient than those with sweep less than the wing. At C L,d , the presence of the LEE planform best satisfying the design criteria was found to have no effect on the wing alone aerodynamic performance.

Leading edge extension↗

NASA NTRS1974

Effects of leading edge sweep angle and design lift coefficient on performance of a modified arrow wing at a design Mach number of 2.6

Wing models were tested in the high-speed section of the Langley Unitary Plan wind tunnel to study the effects of the leading-edge sweep angle and the design lift coefficient on aerodynamic performance and efficiency. The models had leading-edge sweep angles of 69.44 deg, 72.65 deg, and 75.96 deg which correspond to values of the design Mach-number-sweep-angle parameter (beta cotangent A) sub DES of 0.6, 0.75, and 0.9, respectively. For each sweep angle, camber surfaces having design lift coefficients of 0,0.08, and 0.12 at a design Mach number of 2.6 were generated. The wind-tunnel tests were conducted at Mach numbers of 2.3, 2.6, and 2.96 with a stagnation temperature of 338.7 K (150 F) and a Reynolds number per meter of 9.843 times 10 to the 6th power. The results of the tests showed that only a moderate sweeping of the wing leading edge aft of the Mach line along with a small-to-moderate amount of camber and twist was needed to significantly improve the zero-lift (flat camber surface) wing performance and efficiency.

Mack, R. J.↗

NASA NTRS1988

The influence of subsonic mission segments on the use of variable-sweep wings for high speed civil transport configurations

A Mach-3.0, 250-passenger, 6500-n. mi. range SST configuration's alternative use of fixed-planform or variable-sweep wings is presently evaluated, with a view to effects on aerodynamics, mission performance, and sizing. After preliminary design, the fixed and variable-wing configurations were resized to perform missions incorporating subsonic cruise segments of as much as 4000 n. mi.; the effect of subsonic segment length on design gross weight and block time was then ascertained. Due to the reduced supersonic efficiency of the variable-sweep aircraft, over one-half of the 6500-n. mi. mission would have to be flown subsonically for its sizing to reach a lower ramp weight than that of its fixed-geometry counterpart.

Martin, Glenn L.↗

NASA NTRS2016

An Overview of Active Flow Control Enhanced Vertical Tail Technology Development

This paper summarizes a joint NASA/Boeing research effort to advance Active Flow Control (AFC) technology to enhance aerodynamic efficiency of a vertical tail. Sweeping jet AFC technology was successfully tested on subscale and full-scale models as well as in flight. The subscale test was performed at Caltech on a ~14% scale model. More than 50% side force enhancement was achieved by the sweeping jet actuation when the momentum coefficient was 1.7%. AFC caused significant increases in suction pressure on the actuator side and associated side force enhancement. Subsequently, a full-scale Boeing 757 vertical tail model equipped with sweeping jets was tested at the National Full-Scale Aerodynamics Complex 40- by 80-Foot Wind Tunnel at NASA Ames Research Center. There, flow separation control optimization was performed at near flight conditions. Greater than 20% increase in side force were achieved for the maximum rudder deflection of 30deg at the key sideslip angles (0deg and -7.5deg) with a 31-actuator AFC configuration. Based on these tests, the momentum coefficient is shown to be a necessary, but not sufficient parameter to use for design and scaling of sweeping jet AFC from subscale tests to full-scale applications. Leveraging the knowledge gained from the wind tunnel tests, the AFC-enhanced vertical tail technology was successfully flown on the Boeing 757 ecoDemonstrator in the spring of 2015.

Lin, John C.↗

NASA NTRS2010

Experimental Evaluation of Stagnation Point Collection Efficiency of the NACA 0012 Swept Wing Tip

This paper presents the experimental work of a number of icing tests conducted in the Icing Research Tunnel at NASA Glenn Research Center to develop a test method for measuring the local collection efficiency of an impinging cloud at the leading edge of a NACA 0012 swept wing and with the data obtained to further calibrate a proposed correlation for such impingement efficiency calculation as a function of the modified inertia parameter and the sweep angle. The preliminary results showed that there could be some limitation of the test method due to the ice erosion problem when encountered, and also found that, for conditions free of such problem, the stagnation point collection efficiency measurement for sweep angles up to 45 could be well approximated by the proposed correlation. Further evaluation of this correlation is recommended in order to assess its applicability for swept-wing icing scaling analysis.

Tsao, Jen-Ching↗

NASA NTRS1982

Summary and recent results from the NASA advanced High Speed Propeller Research Program

Advanced high-speed propellers offer large performance improvements for aircraft that cruise in the Mach 0.7 to 0.8 speed regime. The current status of the NASA research program on high-speed propeller aerodynamics, acoustics, and aeroelastics is described. Recent wind tunnel results for five 8- to 10-blade advanced models are compared with analytical predictions. Test results show that blade sweep was important in achieving net efficiencies near 80 percent at Mach 0.8 and reducing near-field cruise noise by dB. Lifting line and lifting surface aerodynamic analysis codes are under development and some initial lifting line results are compared with propeller force and probe data. Some initial laser velocimeter measurements of the flow field velocities of an 8-bladed 45 deg swept propeller are shown. Experimental aeroelastic results indicate that cascade effects and blade sweep strongly affect propeller aeroelastic characteristics. Comparisons of propeller near-field noise data with linear acoustic theory indicate that the theory adequate predicts near-field noise for subsonic tip speeds but overpredicts the noise for supersonic tip speeds. Potential large gains in propeller efficiency of 7 to 11 percent at Mach 0.8 may be possible with advanced counter-rotation propellers.

Mitchell, G. A.↗

NASA NTRS1987

Pressure distributions and oil-flow patterns for a swept circulation-control wing

Pressure distributions and photographs of oil flow patterns are presented for a circulation control wing. The model was an aspect ratio four semispan wing mounted on the side wall of the NASA Ames Transonic Wind Tunnel. The airfoil was a 20 percent thick ellipse, modified with circular leading and trailing edges of 4 percent radius, and had a 25.4 cm constant chord. This configuration does not represent a specific wing design, but is generic. A full span, tangetial, rearward blowing, circulation control slot was incorporated ahead of the trailing edge on the upper surface. The wing was tested at Mach numbers from 0.3 to 0.75 at sweep angle of 0 to 45 deg with internal to external pressure ratios of 1.0 to 3.0. Lift and pitching momemt coefficients were obtained from measured pressure distributions at five span stations. When the conventional corrections resulting from sweep angle are applied to the lift and moment of circulation control sections, no additional corrections are necessary to account for changes in blowing efficiency. This is demonstrated for an aft sweep angle of 45 deg. An empirical technique for estimating the downwash distribution of a swept wing was validated.

Keener, Earl R.↗

NASA NTRS1984

Summary of recent NASA propeller research

Advanced high-speed propellers offer large performance improvements for aircraft that cruise in the Mach 0.7 to 0.8 speed regime. At these speeds, studies indicate that there is a 15 to near 40 percent block fuel savings and associated operating cost benefits for advanced turboprops compared to equivalent technology turbofan powered aircraft. Recent wind tunnel results for five eight to ten blade advanced models are compared with analytical predictions. Test results show that blade sweep was important in achieving net efficiencies near 80 percent at Mach 0.8 and reducing nearfield cruise noise by about 6 dB. Lifting line and lifting surface aerodynamic analysis codes are under development and some results are compared with propeller force and probe data. Also, analytical predictions are compared with some initial laser velocimeter measurements of the flow field velocities of an eightbladed 45 swept propeller. Experimental aeroelastic results indicate that cascade effects and blade sweep strongly affect propeller aeroelastic characteristics. Comparisons of propeller near-field noise data with linear acoustic theory indicate that the theory adequately predicts near-field noise for subsonic tip speeds but overpredicts the noise for supersonic tip speeds.

Mikkelson, D. C.↗