Wind-tunnel test of a full-scale, 1.1 pressure ratio, ducted lift-cruise fan
Fixed blade angle, ducted, tip-turbine-driven, cruise fan as aircraft propulsive device at high subsonic speed
SUBSONIC SPEED↗
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Selected winglet and mixed flow long duct nacelle development for DC-10 derivative aircraft
The high speed cruise drag effects of the installation of winglets and a wing tip extension and a mixed flow long duct nacelle are investigated. The winglet program utilized a 4.7 percent semispan model in an eight foot transonic wind tunnel. Winglets provided approximately twice the cruise drag reduction of wing tip extensions for about the same increase in bending moment at the wing-fuselage juncture. The long duct nacelle interference drag program utilized the same model, without the winglets, in the 11 foot transonic wind tunnel. The long duct nacelle, installed in the same position as the current short duct nacelle and with the current production symmetric pylon, was a relatively low risk installation. A pylon with an addition small rearward fairing was also tested and showed some drag reduction potential over the current pylon.
Taylor, A. B.↗
Measuring Electrostatic Discharge
Apparatus measures electrostatic-discharge properties of several materials at once. Allows samples charged either by friction or by exposure to corona. By testing several samples simultaneously, apparatus eliminates errors introduced by variations among test conditions. Samples spaced so they pass at intervals under either of two retractable arms. Samples are 2 inches wide along circular path. Arm tips and voltmeter probe are 6 inches from turntable center. Servocontrolled turntable speed constant within 0.1 percent.
Smith, William C.↗
Advanced Gas Turbine (AGT) Technology Project
Engine testing, ceramic component fabrication and evaluation, component performance rig testing, and analytical studies comprised AGT 100 activities during the 1985 year. Ten experimental assemblies (builds) were evaluated using two engines. Accrued operating time was 120 hr of burning and 170 hr total, bringing cumulative total operating time to 395 hr, all devoid of major failures. Tests identified the generator seals as the primary working fluid leakage sources. Power transfer clutch operation was demonstrated. An alpha SiC gasifier rotor engine test resulted in blade tip failures. Recurring case vibration and shaft whip have limited gasifier shaft speeds to 84%. Ceramic components successfully engine tested now include the SiC scroll assembly, Si3N3 turbine rotor, combustor assembly, regenerator disk bulkhead, turbine vanes, piston rings, and couplings. A compressor shroud design change to reduce heat recirculation back to the inlet was executed. Ceramic components activity continues to focus on the development of state-of-the-art material strength characteristics in full-scale engine hardware. Fiber reinforced glass-ceramic composite turbine (inner) backplates were fabricated by Corning Glass Works. The BMAS/III material performed well in engine testing. Backplates of MAS material have not been engine tested.
Source record↗
Performance of two 10-lb/sec centrifugal compressors with different blade and shroud thicknesses operating over a range of Reynolds numbers
Centrifugal compressors often cannot be directly scaled to very small flow sizes because of structural and manufacturing limitations. The inability to directly scale all design parameters leads to a performance loss other than that which can be associated with the lower Reynolds number. A 10-lb/sec centrifugal compressor was scaled down to 2-lb/sec where adjustments to blade and shroud thickness and fillet radii were required. The modified 2-lb/sec compressor was then directly scaled back up to 10 lb/sec so that the effect of the modifications could be determined. The performance of the two 10-lb/sec compressors is compared over a range of speed and mass flow. The effect of variations in Reynolds number, impeller tip clearance, and shroud thickness on compressor performance is also presented.
Skoch, Gary J.↗
Performance of two 10-lb/sec centrifugal compressors with different blade and shroud thicknesses operating over a range of Reynolds numbers
Centrifugal compressors often cannot be directly scaled to very small flow sizes because of structural and manufacturing limitations. The inability to directly scale all design parameters leads to a performance loss other than that which can be associated with the lower Reynolds number. A 10-lb/sec centrifugal compressor was scaled down to 2-lb/sec where adjustments to blade and shroud thickness and fillet radii were required. The modified 2-lb/sec compressor was then directly scaled back up to 10 lb/sec so that the effect of the modifications could be determined. The performance of the two 10-lb/sec compressors is compared over a range of speed and mass flow. The effect of variations in Reynolds number, impeller tip clearance, and shroud thickness on compressor performance is also presented.
Skoch, Gary J.↗
Noncontact Measurement Of Sizes And Eccentricities Of Holes
Semiautomatic eddy-current-probe apparatus makes noncontact measurements of nominally round holes in electrically conductive specimens and processes measurement data into diameters and eccentricities of holes. Includes x-y translation platform, which holds specimen and moves it horizontally. Probe mounted on probe scanner, positioning probe along vertical (z) direction and rotates probe about vertical axis at preset low speed. Eddy-current sensing coil mounted in side of probe near tip. As probe rotates, impedance analyzer measures electrical impedance (Z) of coil as function of instantaneous rotation angle. Translation and rotation mechanisms and impedance analyzer controlled by computer, which also processes impedance-measurement data.
Chern, Engmin J.↗
Unstructured adaptive mesh computations of rotorcraft high-speed impulsive noise
A new method is developed for modeling helicopter high-speed impulsive (HSI) noise. The aerodynamics and acoustics near the rotor blade tip are computed by solving the Euler equations on an unstructured grid. A stationary Kirchhoff surface integral is then used to propagate these acoustic signals to the far field. The near-field Euler solver uses a solution-adaptive grid scheme to improve the resolution of the acoustic signal. Grid points are locally added and/or deleted from the mesh at each adaptive step. An important part of this procedure is the choice of an appropriate error indicator. The error indicator is computed from the flow field solution and determines the regions for mesh coarsening and refinement. Computed results for HSI noise compare favorably with experimental data for three different hovering rotor cases.
Strawn, Roger↗
Full-Span Tiltrotor Aeroacoustic Model (TRAM) Overview and 40- by 80-Foot Wind Tunnel Test
Most helicopter data trends cannot be extrapolated to tiltrotors because blade geometry and aerodynamic behavior, as well as rotor and fuselage interactions, are significantly different for tiltrotors. A tiltrotor model has been developed to investigate the aeromechanics of tiltrotors, to develop a comprehensive database for validating tiltrotor analyses, and to provide a research platform for supporting future tiltrotor designs. The Full-Span Tiltrotor Aeroacoustic Model (FS TRAM) is a dual-rotor, powered aircraft model with extensive instrumentation for measurement of structural and aerodynamic loads. This paper will present the Full-Span TRAM test capabilities and the first set of data obtained during a 40- by 80-Foot Wind Tunnel test conducted in late 2000 at NASA Ames Research Center. The Full-Span TRAM is a quarter-scale representation of the V-22 Osprey aircraft, and a heavily instrumented NASA and U.S. Army wind tunnel test stand. Rotor structural loads are monitored and recorded for safety-of-flight and for information on blade loads and dynamics. Left and right rotor balance and fuselage balance loads are monitored for safety-of-flight and for measurement of vehicle and rotor aerodynamic performance. Static pressure taps on the left wing are used to determine rotor/wing interactional effects and rotor blade dynamic pressures measure blade airloads. All of these measurement capabilities make the FS TRAM test stand a unique and valuable asset for validation of computational codes and to aid in future tiltrotor designs. The Full-Span TRAM was tested in the NASA Ames Research Center 40- by 80-Foot Wind Tunnel from October through December 2000. Rotor and vehicle performance measurements were acquired in addition to wing pressures, rotor acoustics, and Laser Light Sheet (LLS) flow visualization data. Hover, forward flight, and airframe (rotors off) aerodynamic runs were performed. Helicopter-mode data were acquired during angle of attack and thrust sweeps for a variety of tunnel speeds. Wake geometry images were acquired using LLS photographs and suggest dual tip vortex formation at low thrust conditions. The full paper will include comparisons to isolated-rotor TRAM data acquired at the Duits-Nederlandse Windtunnel (DNW) in 1998. The FS TRAM has been established as a valuable national asset for tiltrotor research. Data reduction and analysis of the 40- by 80-Foot Wind Tunnel test results are underway. Follow-on testing of the FS TRAM is currently being planned for the NASA Ames 80- by 120-Foot Wind Tunnel in late 2001.
McCluer, Megan S.↗
Wind Tunnel Testing of a 120th Scale Large Civil Tilt-Rotor Model in Airplane and Helicopter Modes
In April 2012 and October 2013, NASA and the U.S. Army jointly conducted a wind tunnel test program examining two notional large tilt rotor designs: NASA's Large Civil Tilt Rotor and the Army's High Efficiency Tilt Rotor. The approximately 6%-scale airframe models (unpowered) were tested without rotors in the U.S. Army 7- by 10-foot wind tunnel at NASA Ames Research Center. Measurements of all six forces and moments acting on the airframe were taken using the wind tunnel scale system. In addition to force and moment measurements, flow visualization using tufts, infrared thermography and oil flow were used to identify flow trajectories, boundary layer transition and areas of flow separation. The purpose of this test was to collect data for the validation of computational fluid dynamics tools, for the development of flight dynamics simulation models, and to validate performance predictions made during conceptual design. This paper focuses on the results for the Large Civil Tilt Rotor model in an airplane mode configuration up to 200 knots of wind tunnel speed. Results are presented with the full airframe model with various wing tip and nacelle configurations, and for a wing-only case also with various wing tip and nacelle configurations. Key results show that the addition of a wing extension outboard of the nacelles produces a significant increase in the lift-to-drag ratio, and interestingly decreases the drag compared to the case where the wing extension is not present. The drag decrease is likely due to complex aerodynamic interactions between the nacelle and wing extension that results in a significant drag benefit.
Wind Tunnel Testing of a 1/20th Scale↗
Effect of Trailing Edge Flow Injection on Fan Noise and Aerodynamic Performance
An experimental investigation using trailing edge blowing for reducing fan rotor/guide vane wake interaction noise was completed in the NASA Glenn 9- by 15-foot Low Speed Wind Tunnel. Data were acquired to measure noise, aerodynamic performance, and flow features for a 22" tip diameter fan representative of modern turbofan technology. The fan was designed to use trailing edge blowing to reduce the fan blade wake momentum deficit. The test objective was to quantify noise reductions, measure impacts on fan aerodynamic performance, and document the flow field using hot-film anemometry. Measurements concentrated on approach, cutback, and takeoff rotational speeds as those are the primary conditions of acoustic interest. Data are presented for a 2% (relative to overall fan flow) trailing edge injection rate and show a 2 dB reduction in Overall Sound Power Level (OAPWL) at all fan test speeds. The reduction in broadband noise is nearly constant and is approximately 1.5 dB up to 20 kHz at all fan speeds. Measurements of tone noise show significant variation, as evidenced by reductions of up to 6 dB in the 2 BPF tone at 6700 rpm.: and increases of nearly 2 dB for the 4 BPF tone at approach speed. Aerodynamic performance measurements show the fan with 2 % injection has an overall efficiency that is comparable to the baseline fan and operates, as intended, with nearly the same pressure ratio and mass flow parameters. Hot-film measurements obtained at the approach operating condition indicate that mean blade wake filling in the tip region was not as significant as expected. This suggests that additional acoustic benefits could be realized if the trailing edge blowing could be modified to provide better filling of the wake momentum deficit. Nevertheless, the hot-film measurements indicate that the trailing edge blowing provided significant reductions in blade wake turbulence. Overall, these results indicate that further work may be required to fully understand the proper implementation of injecting flow at/near the trailing edge as a wake filling strategy. However, data do support the notion that noise reductions can be realized not only for tones but perhaps more importantly, also for broadband. Furthermore, the technique can be implemented without adversely effecting overall fan aerodynamic performance.
Fite, E. Brian↗
Upper Surface Nacelle Influence on SCAR Aerodynamic Characteristics at Transonic Speeds
The arrow-wing transport configuration with detached engines located over the wing to produce upper surface exhaust flow effects was tested at angles of attack from -4 deg to 8 deg and jet total-pressure ratios from 1 (Jet off) to approximately 10. Wing tip leading edge flap deflections of -10 deg to 10 deg were tested with the wing-body configuration only (no nacelles). Tests were made with various nacelle chordwise, spanwise, and vertical height locations over the Mach number, angle of attack, and jet total-pressure ratio ranges. Deflecting the wing tip leading edge flap from 0 deg to -10 deg increased maximum lift to drag ratio by 1.0 at subsonic speeds. Installation of upper surface nacelles (no wing/nacelle pylons) increased the wing-body pitching moment at all Mach numbers and decreased the drag of the wing-body configuration at subsonic Mach numbers. Jet exhaust interference effects were negligible.
Mercer, C. E.↗
A Navier-Stokes study of shock-boundary layer interaction and flow separation inside a transonic compressor
A numerical study to evaluate a three-dimensional Navier-Stokes method as a tool to predict the detailed flowfield inside a low-aspect-ratio compressor at off-design rotor speed has been conducted. The flow field inside a state-of-the-art transonic compressor is used for the purpose of the evaluation. The experimental study shows that the rotor has higher peak efficiency at 90 percent rotor speed than at 100 percent rotor speed. The details of the flow structure inside the low-aspect-ratio compressor (three-dimensional shock structure, shock-boundary layer interaction, tip leakage vortex, etc.) and the overall aerodynamic performance at various operating conditions are numerically analyzed, and the results are compared with the available experimental data. The numerical results also indicate that the rotor has higher peak efficiency at 90 percent rotor speed than at 100 percent rotor speed. This is due to the reduced shock strength and the reduced interactions among passage shock, tip-clearance vortex and blade boundary layer.
Hah, Chunill↗
An experimental investigation of wing tip turbulence with applications to aerosound
Wind tunnel tests were performed to measure the turbulence characteristics of flow in the wing tip region of a blunt tipped NACA 0012 airfoil at angles of attack of 16, 12 and 6 deg and flow speeds of 75, 55, and 35 m/s. Reynolds numbers based on the chord length ranged from 1.8 x 10(exp6) to 3.8 x 10(exp 6). Hot wire measurements were made at an angle of attack of 12 deg and U(infinity) = 55 ms. These indicate fluctuation velocities on the order of 0.13 U(infinity) in the tip region. Surface pressure spectra are presented for alpha = 12 deg and 16 deg; surface-pressure cross-correlations are presented for alpha = 16 deg. These suggest the presence of a well organized turbulence structure with a peak Strouhal number 0.8 to 1.1 based on the wing thickness. High frequency turbulence of lesser intensity was measured at positions near the primary tip vortex on the upper wing surface. A separation region with a spectral peak of St approximately equal to 0.4 was recorded at alpha = 16 deg for the transducer located on the low pressure side of the wing nearest the trailing edge. Surface-surface cross correlations display a double, echo like structure. Near-far field correlations indicate that the well ordered tip turbulence and the separation region on the upper surface both radiate significant levels of far field sound. The use of cross correlation techniques in the wind tunnel has permitted treatment of radiated sound levels which are 25 and more dB below tunnel background levels.
Mcinerny, S. A.↗
Clearance Sensitivity Mitigation in Small Core Compressors
As future aircraft become lighter and more aerodynamically efficient, thrust requirements will decrease, reducing the core size of the engine. Furthermore, in the pursuit of improved fuel burn, engine overall pressure ratio and bypass ratio will increase, further driving down engine core size. These drivers together mean that the core size, or corrected mass flow rate at the compressor exit, for future single-aisle aircraft applications will shrink below 3.0 lbm/s. Traditionally, this small core compressor size is in the domain of axi-centrifugal designs, machines that are typically less efficient and limited to pressure ratios of ~25 due to stress and thermomechanical fatigue in the centrifugal impeller. In this light, NASA, Pratt & Whitney and the Raytheon Technologies Research Center embarked upon a program to develop technologies to enable an all-axial high-pressure compressor with a core size below 3.0 lbm/s and an overall pressure ratio greater than 50. One of the challenges with an all-axial high-pressure compressor at this core size is the small span at the rear of the compressor. As core size is scaled down, the rotor tip clearances do not scale with span, leading to significant efficiency penalties. This paper documents a numerical and experimental effort to mitigate this penalty through design optimization and technology insertion, enabling an estimated 5 to 10% fuel burn reduction relative to 2020 best-in-class. The experimental program consisted of two builds of a high-speed rig: a baseline build and a second build to insert technology to manage large rotor tip gap/span ratios. The results demonstrated a reduction in the sensitivity of the compressor to large clearance ratios, which would be likely at the end-of-life of a small core application. The test campaign demonstrated that, with the insertion of technology, small core compressors can maintain a similar efficiency to current best-in-class large core size compressors. In addition to advancing the state-of-the-art of technology, the program also advanced the modeling standards for multistage compressors with large clearance-to-span ratios. The validation of this approach is described in the paper.
Axial Compressors↗
Clearance Sensitivity Mitigation in Small Core Compressors
As future aircraft become lighter and more aerodynamically efficient, thrust requirements will decrease, reducing the core size of the engine. Furthermore, in the pursuit of improved fuel burn, engine overall pressure ratio and bypass ratio will increase, further driving down engine core size. These drivers together mean that the core size, or corrected mass flow rate at the compressor exit, for future single-aisle aircraft applications will shrink below 3.0 lb/s. Traditionally, this small core compressor size is in the domain of axi-centrifugal designs, machines that are typically less efficient and limited to pressure ratios of ~25 due to stress and thermomechanical fatigue in the centrifugal impeller. In this light, NASA, Pratt & Whitney and the Raytheon Technologies Research Center embarked upon a program to develop technologies to enable an all-axial high-pressure compressor with a core size below 3.0 lb/s and an overall pressure ratio greater than 50. One of the challenges with an all-axial high-pressure compressor at this core size is the small span at the rear of the compressor. As core size is scaled down, the rotor tip clearances do not scale with span, leading to significant efficiency penalties. This paper documents a numerical and experimental effort to mitigate this penalty through design optimization and technology insertion, enabling an estimated 5 to 10% fuel burn reduction relative to 2020 best-in-class. The experimental program consisted of two builds of a high-speed rig: a baseline build and a second build to insert technology to manage large rotor tip gap/span ratios. The results demonstrated a reduction in the sensitivity of the compressor to large clearance ratios, which would be likely at the end-of-life of a small core application. The test campaign demonstrated that, with the insertion of technology, small core compressors can maintain a similar efficiency to current best-in-class large core size compressors. In addition to advancing the state-of-the-art of technology, the program also advanced the modeling standards for multistage compressors with large clearance-to-span ratios. The validation of this approach is described in the paper.
Axial Compressors↗
Performance, Inflow, and Tip Loss Characteristics of Rotors with Discontinuous Steps in Twist and Lift at the Blade-Tips
This paper examines the inflow and tip-loss characteristics of rotors incorporating indexed blade-tips that have discontinuous steps in twist or lift distribution. An extension of the classic two-dimensional Prandtl tip loss analysis is derived for indexed-tip rotors and proprotors. Additionally, parametric analysis of tip span and indexed-tip twist-steps was performed using computational fluid dynamics for hover and high-speed axial-flow conditions for proprotors. The intent of this work is to define a new class of active-rotor-twist-control designs for proprotors that avoids unnecessary aeroperformance compromises for rotor twist distributions for proprotors in all phases of flight.
rotors↗
Formulas for the Supersonic Loading, Lift, and Drag of Flat Swept-Back Wings with Leading Edges Behind the Mach Line
The method of superposition of linearized conical flows has been applied to the calculation of the aerodynamic properties, in supersonic flight, of thin flat, swept-back wings at an angle of attack. The wings are assumed to have rectilinear plan forms, with tips parallel to the stream, and to taper in the conventional sense. The investigation covers the moderately supersonic speed range where the Mach lines from the leading-edge apex lie ahead of the wing. The trailing edge may lie ahead of or behind the Mach lines from its apex. The case in which the Mach cone from one tip intersects the other tip is not treated. Formulas are obtained for the load distribution, the total lift, and the drag due to lift. For the cases in which the trailing edge is outside the Mach cone from its apex the formulas are complete. For wings with both leading and trailing edges behind their respective Mach lines, a degree of approximation is necessary. Charts of some of the functions derived are included to facilitate computing, and several examples are worked out in outline.
Cohen, Doris↗