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At least 235 records · Page 13

Functional Task Tests in Partial Gravity During Parabolic Flight

BACKGROUND Critical mission tasks required by crews immediately after landing on a planetary surface include walking, jumping, and egressing from a seat. Understanding how these functional tasks are performed in partial gravity such as on the moon or Mars is necessary to define effective and comprehensive countermeasure strategies for preserving crew performance during exploration-class missions. We propose to study the performance of these tasks during the partial gravity phases of parabolic flight. These tasks will be performed using the same equipment and procedures as those used with astronauts returning from spaceflight and with ground-based subjects after prolonged axial body unloading during bed rest (sensorimotor standard measures). HYPOTHESIS We hypothesize that partial gravity during parabolic flight will cause acute changes in vestibular, proprioceptive, and sensorimotor functions, and these changes will impact the performance of mission critical tasks such as standing, walking, and jumping. The largest changes in performance are expected at the lowest gravity level (0.25g) because subjects will no longer be able to use the gravitational reference for the perception of upright. Ultimately, this information could be used to assess performance risks and inform the design of countermeasures for NASA exploration-class human missions. METHODS Twelve subjects will be tested during three flights of 30 parabolas, including 10 parabolas at 0.25g, 10 parabolas at 0.5g, and 10 parabolas at 0.75g. Subjects also will perform tests in 1g between parabolas. The tasks will be the same as those tested on astronauts returning from spaceflight: sit-to-stand with obstacle walk, tandem rail balance, jump down, and recovery from fall. Measurements will include: (a) time to test completion (sit-to-stand with obstacle walk, recovery from fall); (b) time elapsed between the start of motion and the stabilization of upright posture (recovery from fall, jump down); (c) mean sway speed during quiet standing (recovery from fall, jump down); (d) changes in heart rate and blood pressure (recovery from fall); (e) balance time and torso accelerations (tandem rail balance); (f) cone of stability (jump down); and (g) severity of motion sickness symptoms. RELEVANCE Although gravitational dose-response curves have been obtained for some biochemical systems in animals, these dose-responses are unknown for most human physiologic systems. Our study will compare the outcomes of four functional task tests in 0.25g, 0.5g, 0.75g, and 1g with those previously obtained in ground-based subjects after prolonged axial body unloading and in astronauts immediately after spaceflight. These comparisons will help us understand the true extent of functional task performance deficits in partial gravity. The dose-response relationship between gravity level and task performance decrement also will help determining the gravity threshold for these functional tasks. ACKNOWLEDGEMENT This work is supported by the NASA’s Human Research Program Human Health Countermeasures Element.

T. R. Macaulay↗

Flight, Wind Tunnel, and Numerical Experiments with a Slender Cone at Incidence

The three-dimensional leeward separation about a 5-deg semi-angle cone of 11 deg angle of incidence was Investigated in night, in the wind tunnel, and by numerical computations. The test conditions were Mach numbers of 0.6, 1.5, and 1.8 at Reynolds numbers between 7 and 10 million based on freestream conditions and a 76.2-cm (30-in.) length of surface. The surface pressure conditions measured included those of fluctuating and mean static, as well as recovery pressures generated by obstacle blocks to provide skin friction and separation-line locations. The mean static pressures from flight and wind tunnel were in reasonably good agreement. The computed results gave the same distributions, but were slightly more positive in magnitude. The experimentally measured primary and secondary separation line locations compared closely with computed results. There were substantial differences In level between the surface root-mean-square pressure fluctuations obtained in night and in the wind tunnel, due, It Is thought, to a relatively high acoustic disturbance level in the tunnel compared with the quiescent atmospheric conditions in night.

Peake, David J.↗

Balanced Flow Metering and Conditioning: Technology for Fluid Systems

Revolutionary new technology that creates balanced conditions across the face of a multi-hole orifice plate has been developed, patented and exclusively licensed for commercialization. This balanced flow technology simultaneously measures mass flow rate, volumetric flow rate, and fluid density with little or no straight pipe run requirements. Initially, the balanced plate was a drop in replacement for a traditional orifice plate, but testing revealed substantially better performance as compared to the orifice plate such as, 10 times better accuracy, 2 times faster (shorter distance) pressure recovery, 15 times less acoustic noise energy generation, and 2.5 times less permanent pressure loss. During 2004 testing at MSFC, testing revealed several configurations of the balanced flow meter that match the accuracy of Venturi meters while having only slightly more permanent pressure loss. However, the balanced meter only requires a 0.25 inch plate and has no upstream or downstream straight pipe requirements. As a fluid conditioning device, the fluid usually reaches fully developed flow within 1 pipe diameter of the balanced conditioning plate. This paper will describe the basic balanced flow metering technology, provide performance details generated by testing to date and provide implementation details along with calculations required for differing degrees of flow metering accuracy.

Kelley, Anthony R.↗

Performance of vortex generators in a Mach 2.5 low-bleed full scale 45-percent-internal-contraction axisymmetric inlet

Steady-state and dynamic flow characteristics associated with two sets of vortex generators having different mixing criteria were determined. The inlet performance with and without these vortex generators is presented. The vortex generators were successful in eliminating separation, increasing area-weighted total pressure recovery, and decreasing distortion. Transmission times obtained from cross-correlations of the wall static pressures and the diffuser exit total pressure showed no effect of the upstream flow characteristics on the diffuser exit pressures when generators were used. Without generators, separation occurred and the upstream pressure characteristics had immediate effects on the diffuser exit pressure characteristics.

Neumann, H. E.↗

Investigation of Inlet Concepts for Maneuver Improvement at Transonic Speeds

A 15 percent scale lightweight fighter type inlet forebody was tested in the Ames 14 foot transonic wind tunnel at Mach numbers of 0.7, 0.9, and 1.04. The inlet was a two dimensional horizontal ramp system designed for a Mach number of 2.2. Four inlet devices designed to prevent or delay cowl-lip boundary layer separation or to improve the inlet internal flow characteristics at high angles of attack were investigated. The devices used to control cowl-lip separation consisted of cowl leading edge flaps, slotted flaps, and tangential blowing. To improve the internal flow characteristics, discrete jet nozzle flows were directed downstream and parallel to the duct surface in the subsonic diffuser to energize the wall boundary layer. The discrete jets used in the subsonic diffuser were also tested in combination with each of the cowl leading edge devices. Test measurements included engine-face total pressure recovery, steady state distortion, dynamic distortion, duct boundary layer profiles, and duct-surface static pressures.

Latham, E.↗

Experimental investigation of a 0.15-scale model of an underfuselage normal-shock inlet

A 0.15 scale model of an underfuselage inlet designed for a single-engine fighter airplane was tested. The inlet was a fixed-geometry, normal-shock configuration designed to operate at flight speeds up to Mach 2.0. Peformance data for the basic inlet and several configuration variations are presented as a function of angle of attack, angle of sideslip, and airflow in the 0.6 to 2.0 Mach number range. The configuration variations included boundary-layer diverter height, cowl and splitter-plate modifications, and inlet bleed system variations. Flow-field characteristics at the simulated engine face, at the inlet throat, at the splitter-plate leading edge, and forward of the inlet are presented. The pressure recovery of the inlet is approximately equal to the product of theoretical normal-shock and duct pressure recoverable at cruise angle of attack. Very good performance at high angle of attack was obtained. Pressure distortion and turbulence at the engine face were low, and the inlet remained stable at all engine airflows over the flight maneuver envelope of the aircraft for which the inlet was designed.

Leamer, P. C.↗

High-pressure phase transition in 3-D printed nanolamellar high-entropy alloy by imaging and simulation insights

Abstract We report on the high-resolution imaging and molecular dynamics simulations of a 3D-printed eutectic high-entropy alloy (EHEA) Ni 40 Co 20 Fe 10 Cr 10 Al 18 W 2 consisting of nanolamellar BCC and FCC phases. The direct lattice imaging of 3D-printed samples shows the Kurdjumov–Sachs (K–S) orientation relation {111} FCC parallel to {110} BCC planes in the dual-phase lamellae. Unlike traditional iron and steels, this alloy shows an irreversible BCC-to-FCC phase transformation under high pressures. The nanolamellar morphology is maintained after pressure cycling to 30 GPa, and nano-diffraction studies show both layers to be in the FCC phase. The chemical compositions of the dual-phase lamellae after pressure recovery remain unchanged, suggesting a diffusion-less BCC–FCC transformation in this EHEA. The lattice imaging of the pressure-recovered sample does not show any specific orientation relation between the two resulting FCC phases, indicating that many grain orientations are produced during the BCC–FCC phase transformation. Molecular dynamics simulations on phase transformation in a nanolamellar BCC/FCC in K–S orientation show that phase transformation from BCC to FCC is completed under high pressures, and the FCC phase is retained on decompression aided by the stable interfaces. Our work elucidates the irreversible phase transformation under static compression, providing an understanding of the orientation relationships in 3-D printed EHEA under high pressures.

3-D Printing↗

Assessment of Using Ideal Gas for Predicting Boattail Flow at Cryogenic Temperatures

The applicability of using ideal gas assumptions to simulate high Reynolds number experimental data that was obtained at cryogenic temperatures is examined. Flow over an axisymmetric nozzle boattail model was calculated using reference temperatures of 117 K and 300 K and at Reynolds numbers from 50 to 200 million per meter. From the testing perspective, pressure, compression factor, and isentropic coefficients calculated using one-dimensional real gas equations are used to examine the departure of cryogenic flow from ideal gas flow across the range of temperatures and potential impacts on measured aerodynamic data. Solutions developed using ideal gas assumptions in a three-dimensional Navier-Stokes code are compared with experimental data obtained at cryogenic temperatures at two unit Reynolds numbers at freestream Mach numbers of 0.6 and 0.9. Results for several one- and two-equation turbulence models are shown. Predicted pressure coefficient distributions along the nozzle boattail differed from experimental data between 8% to less than 0.5% depending on the turbulence model and Mach number. The greatest discrepancy occurred in the level of static pressure recovery in the recompression region where the flow was separated. Solutions using warm and cryogenic freestream temperatures predicted similar boattail pressure distributions at the same unit Reynolds number.

Nozzle↗

Assessment of Using Ideal Gas for Predicting Boattail Flow at Cryogenic Temperatures

The applicability of using ideal gas assumptions to simulate high Reynolds number experimental data that was obtained at cryogenic temperatures is examined. Flow over an axisymmetric nozzle boattail model was calculated using reference temperatures of 117 K and 300 K and at Reynolds numbers from 50 to 200 million per meter. From the testing perspective, pressure, compression factor, and isentropic coefficients calculated using one-dimensional real gas equations are used to examine the departure of cryogenic flow from ideal gas flow across the range of temperatures and potential impacts on measured aerodynamic data. Solutions developed using ideal gas assumptions in a three-dimensional Navier-Stokes code are compared with experimental data obtained at cryogenic temperatures at two unit Reynolds numbers at freestream Mach numbers of 0.6 and 0.9. Results for several one- and two-equation turbulence models are shown. Predicted pressure coefficient distributions along the nozzle boattail differed from experimental data between 8% to less than 0.5% depending on the turbulence model and Mach number. The greatest discrepancy occurred in the level of static pressure recovery in the recompression region where the flow was separated. Solutions using warm and cryogenic freestream temperatures predicted similar boattail pressure distributions at the same unit Reynolds number.

Nozzle↗

Hot gas ingestion characteristics and flow visualization of a vectored thrust STOVL concept

A 9.2 percent scale short takeoff and vertical landing (STOVL) hot gas ingestion model was designed and built by McDonnell Douglas Corporation (MCAIR) and tested in the NASA Lewis Research Center 9- by 15-Foot Low Speed Wind Tunnel (LSWT). Hot gas ingestion, the entrainment of heated engine exhaust into the inlet flow field, is a key development issue for advanced short takeoff and vertical landing aircraft. The Phase 1 test program, conducted by NASA Lewis and McDonnell Douglas Corporation, evaluated the hot ingestion phenomena and control techniques and Phase 2 test program which was conducted by NASA Lewis are both reported. The Phase 2 program was conducted at exhaust nozzles temperatures up to 1460 R and utilized a sheet laser system for flow visualization of the model flow field in and out of ground effects. Hot gas ingestion levels were measured for the several forward nozzle splay configurations and with flow control/lift improvement devices which reduced the hot gas ingestion. The model support system had four degrees of freedom, heated high pressure air for nozzle flow, and a suction system exhaust for inlet flow. The headwind (freestream) velocity for Phase 1 was varied from 8 to 90 kn, with primary data taken in the 8 to 23 kn headwind velocity range. Phase 2 headwind velocity varied from 10 to 23 kn. Results of both Phase 1 and 2 are presented. A description of the model, facility, a new model support system, and a sheet laser illumination system are also provided. Results are presented over a range of main landing gear height (model height) above the ground plane at a 10 kn headwind velocity. The results contain the compressor face pressure and temperature distortions, total pressure recovery, compressor face temperature rise, and the environmental effects of the hot gas. The environmental effects include the ground plane temperature and pressure distributions, model airframe heating, and the location of the ground flow separation. Results from the sheet laser flow visualization test are also shown.

Johns, Albert L.↗

Altitude-Wind-Tunnel Investigation of R-4360-18 Power-Plant Installation for XR60 Airplane: Performance of Induction and Exhaust Systems - 3

A study has been made of the performance of the induction and the exhaust systems on the XR60 power-plant installation as part of an investigation conducted in the Cleveland altitude wind tunnel. Altitude flight conditions from 5000 to 30,000 feet were simulated for a range of engine powers from 750 to 3000 brake horsepower. Slipstream rotation prevented normal pressure recoveries in the right side of the main duct in the region of the right intercooler cooling-air duct inlet. Total-pressure losses in the charge-air flow between the turbosupercharger and the intercoolers were as high as 2.1 inches of mercury. The total-pressure distribution of the charge air at the intercooler inlets was irregular and varied as much as 1.0 inch of mercury from the average value at extreme conditions, Total-pressure surveys at the carburetor top deck showed a variation from the average value of 0.3 inch of mercury at take-off power and 0.05 inch of mercury at maximum cruising power, The carburetor preheater system increased the temperature of the engine charge air a maximum of about 82 F at an average cowl-inlet air temperature of 9 F, a pressure altitude of 5000 feet, and a brake horsepower of 1240.

Dupree, David T.↗

Total-Pressure Distortion and Recovery of Supersonic Nose Inlet with Conical Centerbody in Subsonic Icing Conditions

Ice was formed on a full-scale unheated supersonic nose inlet in the NACA Lewis icing tunnel to determine its effect on compressor-face total-pressure distortion and recovery.Inlet angle of attack was varied from 0degrees to 12 degrees, free-stream Mach number from 0.17 to 0.28, and compressor-face Mach number from 0.10 to 0.47. Icing-cloud liquid-water content was varied from 0.65 to 1.8 grams per cubic meter at free-stream static air temperatures of 15 degrees and 0 degrees F. The addition of ice to the inlet components increased total-pressure-distortion levels and decreased recovery values compared withclear0air results, the losses increasing with time in ice. The combination of glaze ice, high corrected weight flow, and high angle of attack yielded the highest levels of distortion and lowest values of recovery. The general character of compressor-face distortion with an iced inlet was the same as that for the clean inlet, the total-pressure gradients being predominantly radial, with circumferential gradients occurring at angle of attack. At zero angle of attack, free-stream Mach number of 0.27, and a constant corrected weight flow of 150 pounds per second (compressor-face Mach number of 0.43), compressor-face total-pressure-distortion level increased from about 6 percent in clear air to 12 percent after 21 minutes of heavy glaze icing; concurrently, total-pressure recovery decreased from about 0.98 to 0.945. For the same operating conditions but with the inlet at 12 deg angle of attack, a change in distortion level occurred from about 9 percent in clear air to 14 percent after 2-1/4 minutes of icing, with a decrease in recovery from about 0.97 to 0.94.

Gelder, Thomas F↗

Experimental investigation of a Mach 6 fixed-geometry inlet featuring a swept external-internal compression flow field

An investigation of a fixed-geometry, swept external-internal compression inlet was conducted at a Mach number of 6.0 and a test-section Reynolds number of 1.55 x 10 to the 7th power per meter. The test conditions was constant for all runs with stagnation pressure and temperature at 20 atmospheres and 500 K, respectively. Tests were made at angles of attack of -5 deg, 0 deg, 3 deg, and 5 deg. Measurements consisted of pitot- and static-pressure surveys in inlet throat, wall static pressures, and surface temperatures. Boundary-layer bleed was provided on the centerbody and on the cowl internal surface. The inlet performance was consistently high over the range of the angle of attack tested, with an overall average total pressure recovery of 78 percent and corresponding adiabatic kinetic-energy efficiency of 99 percent. The inlet throat flow distribution was uniform and the Mach number and pressure level were of the correct magnitude for efficient combustor design. The utilization of a swept compression field to meet the starting requirements of a fixed-geometry inlet produced neither flow instability nor a tendency to unstart.

Torrence, M. G.↗

Long duct nacelle aerodynamic development for DC-10 derivatives

The results are presented of a wind tunnel test utilizing a 4.7-percent-scale semispan model of the DC-10 in the Calspan 8-foot transonic wind tunnel. The effect of a revised long-duct nacelle shape on the channel velocities, the incremental drag relative to the baseline long-duct nacelle, and channel velocities for the baseline long-duct nacelle were determined and compared with data obtained at Ames. The baseline and the revised long-duct nacelles are representative of a CF6-50 mixed-flow configuration and were evaluated on a model of a proposed DC-10 stretched-fuselage configuration. The results showed that the revised long-duct nacelle has an appreciable effect on the inboard channel velocities, resulting in an increased channel Mach number. However, the pressure recovery on the nacelle afterbody was about the same for both nacelles. The lift curves for both long-duct nacelle configurations were the same. The channel pressures measured at Calspan were in good agreement with those measured at Ames for the baseline long-duct nacelle. The incremental drag for the revised nacelle was measured as two to four counts (three counts is approximately equal to one percent of the airplane drag) higher than that of the baseline long-duct nacelle.

Patel, S. P.↗

Performance and boundary-layer evaluation of a sonic inlet

Tests were conducted to determine the boundary layer characteristics and aerodynamic performance of a radial vane sonic inlet with a length/diameter ratio of 1 for several vane configurations. The sonic inlet was designed with a slight wavy wall type of diffuser geometry, which permits operation at high inlet Mach numbers (sufficiently high for good noise suppression) without boundary layer flow separation and with good total pressure recovery. A new method for evaluating the turbulent boundary layer was developed to separate the boundary layer from the inviscid core flow, which is characterized by a total pressure variation from hub to tip, and to determine the experimental boundary layer parameters.

Schmidt, J. F.↗

Airfoil design for Reynolds numbers between 50,000 and 500,000

The design of airfoils for flows with Re of 50,000-500,000 requires consideration of laminar separation bubbles. A design approach is discussed which specifies the angle of attack at which the potential flow velocity is to be constant at each segment of the airfoil. The velocity gradient is controlled by introducing a pressure recovery function at the trailing edge. Boundary layer stability decreases with rising Re, although an upper Re value can be identified, below which the boundary layer will be stable. Adverse pressure gradients are associated with the shape parameter of the velocity profile, whose rise in value decreases stability. Transition displays similar relationships to the shape parameter. The most frequent feature of separation is the appearance of a separation bubble.

R Eppler↗

Transonic low aspect ratio wing-winglet designs

A numerical design study has been conducted to ascertain the potential of winglets as a drag-reducing measure at high subsonic Mach numbers for low aspect ratio wings. The four variants of the winglet concept studied are a 'detuned' winglet with decreased incidence at the wing-winglet juncture; a steerable winglet; more gradual pressure recovery at the wing and winglet trailing edges; and the application of supercritical airfoil technology. A further study is conducted to assess the accuracy of the numerical code's predicted pressure drag values.

Kuhlman, John M.↗

Flight Tests of the Effect of Several Modifications on the Maximum Speed of the P-63A Airplane

Presented herein are the results of flight tests conducted to obtain the effect on maximum speed of several modifications to the P-63A airplane. These modifications made to the P-63A airplane as a result of previous flight and full-scale-tunnel tests increased the maximum speed of the airplane by 6 miles per hour. About one-half the increase is attributed to a reduction in drag and the remainder to increased ram at the carburetor entrance. The increase in ram of approximately 0.24 qc was obtained through modifications to the carburetor-scoop entrance and duct. The auxiliary supercharger is hydraulically coupled to the engine in such a way that the speed of the supercharger is regulated by absolute carburetor total pressure. This regulation is set so that full coupling is not reached until the airplane is well above full-throttle altitude. The action of the increased ram in reducing the speed of the auxiliary supercharger in the altitude range just above full-throttle altitude prevented a further speed increase of approximately 2 miles per hour from accompanying the increase in ram. Also, if the· auxiliary supercharger control could be rearranged to permit the maximum supercharger pressure ratio to be obtained at the full-throttle altitude, the speed in either the original or the modified configuration would be increased 5 miles per hour above that actually measured. The pressure recovery on the face of the prestone and oil radiators in the high-speed flight conditions was increased by an average of 14 percent of free-stream impact pressure as a result of the modifications made to the wing duct inlets.

Thomas J Voglewede↗