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At least 1,531 records · Page 85
All sourcesNASA NTRSDOE OSTI
NASA NTRS1955

Performance of 15-Stage Experimental J71 Axial-Flow Compressor. III - Effects of Inlet-Guide-Vane Adjustment

The stall-limit line at low speeds was improved somewhat by closing the inlet guide vanes 6 deg, while the design-speed maximum flow and pressure ratio were reduced. The first-stage characteristic curve was moved to lower values of both flog coefficient and equivalent pressure ratio. The second-stage pressure ratio was decreased slightly at high speeds, while the later stages were unaffected.

Lucas, James G.↗
NASA NTRS2009

Improvements in Modeling 90 degree Bleed Holes for Supersonic Inlets

The modeling of porous bleed regions as boundary conditions in computational fluid dynamics (CFD) simulations of supersonic inlet flows has been improved through a scaling of sonic flow coefficient data for 90deg bleed holes. The scaling removed the Mach number as a factor in computing the sonic flow coefficient and allowed the data to be fitted with a quadratic equation, with the only factor being the ratio of the plenum static pressure to the surface static pressure. The implementation of the bleed model into the Wind-US CFD flow solver was simplified by no longer requiring the evaluation of the flow properties at the boundary-layer edge. The quadratic equation can be extrapolated to allow the modeling of small amounts of blowing, which can exist when recirculation of the bleed flow occurs within the bleed region. The improved accuracy of the bleed model was demonstrated through CFD simulations of bleed regions on a flat plate in supersonic flow with and without an impinging oblique shock. The bleed model demonstrated good agreement with experimental data and three-dimensional CFD simulations of bleed holes.

Slater, John W.↗
NASA NTRS2010

Low-Dead-Volume Inlet for Vacuum Chamber

Gas introduction from near-ambient pressures to high vacuum traditionally is accomplished either by multi-stage differential pumping that allows for very rapid response, or by a capillary method that allows for a simple, single-stage introduction, but which often has a delayed response. Another means to introduce the gas sample is to use the multi-stage design with only a single stage. This is accomplished by using a very small conductance limit. The problem with this method is that a small conductance limit will amplify issues associated with dead -volume. As a result, a high -vacuum gas inlet was developed with low dead -volume, allowing the use of a very low conductance limit interface. Gas flows through the ConFlat flange at a relatively high flow rate at orders of magnitude greater than through the conductance limit. The small flow goes through a conductance limit that is a double-sided ConFlat.

Naylor, Guy↗
NASA NTRS2011

The Evolution of Utilizing Manual Throttles to Avoid Low LH2 NPSP at the SSME Inlet

Even before the first flight of the Space Shuttle, it was understood low liquid hydrogen (LH2) Net Positive Suction Pressure (NPSP) at the inlet to the Space Shuttle Main Engine (SSME) can have adverse effects on engine operation. A number of failures within both the External Tank (ET) and the Orbiter Main Propulsion System could result in a low LH2 NPSP condition. Operational workarounds were developed to take advantage of the onboard crew s ability to manually throttle down the SSMEs, which alleviated the low LH2 NPSP condition. A throttling down of the SSME resulted in an increase in NPSP, mainly due to the reduction in frictional flow losses while at a lower throttle setting. As engineers refined their understanding of the NPSP requirements for the SSME (through a robust testing program), the operational techniques evolved to take advantage of these additional capabilities. Currently the procedure, which for early Space Shuttle missions required a Return-to-Launch-Site abort, now would result in a nominal Main Engine Cut Off (MECO) and no loss of mission objectives.

Henfling, Rick↗
NASA NTRS2010

The Evolution of Utilizing Manual Throttling to Avoid Excessively Low LH2 NPSP at the SSME Inlet

In the late 1970s, years before the Space Shuttle flew its maiden voyage, it was understood low liquid hydrogen (LH2) Net Positive Suction Pressure (NPSP) at the inlet to the Space Shuttle Main Engine (SSME) could have adverse effects on engine operation. A number of failures within both the External Tank (ET) and the Orbiter Main Propulsion System (MPS) could result in a low LH2 NPSP condition, which at extremely low levels can result in cavitation of SSME turbomachinery. Operational workarounds were developed to take advantage of the onboard crew s ability to manually throttle down the SSMEs (via the Pilot s Speedbrake/Throttle Controller), which alleviated the low LH2 NPSP condition. Manually throttling the SSME to a lower power level resulted in an increase in NPSP, mainly due to the reduction in frictional flow losses while at the lower throttle setting. Early in the Space Shuttle Program s history, the relevant Flight Rule for the Booster flight controller in Mission Control did not distinguish between ET and Orbiter MPS failures and the same crew action was taken for both. However, after a review of all Booster operational techniques following the Challenger disaster in the late 1980s, it was determined manually throttling the SSME to a lower power was only effective for Orbiter MPS failures and the Flight Rule was updated to reflect this change. The Flight Rule and associated crew actions initially called for a single throttle step to minimum power level when a low threshold for NPSP was met. As engineers refined their understanding of the NPSP requirements for the SSME (through a robust testing program), the operational techniques evolved to take advantage of the additional capabilities. This paper will examine the evolution of the Flight rule and associated procedure and how increases in knowledge about the SSME and the Space Shuttle vehicle as a whole have helped shape their development. What once was a single throttle step when NPSP decreased to a certain low threshold has now become three throttle steps, each occurring at a lower NPSP threshold. Additionally the procedure, which for early Space Shuttle missions required a Return-to-Launch-Site abort, now results in a nominal Main Engine Cut Off and no loss of mission objectives.

Henfling, Rick↗
NASA NTRS2011

Low-Dead-Volume Inlet for Vacuum Chamber

Gas introduction from near-ambient pressures to high vacuum traditionally is accomplished either by multi-stage differential pumping that allows for very rapid response, or by a capillary method that allows for a simple, single-stage introduction, but which often has a delayed response. Another means to introduce the gas sample is to use the multi-stage design with only a single stage. This is accomplished by using a very small conductance limit. The problem with this method is that a small conductance limit will amplify issues associated with dead-volume. As a result, a high-vacuum gas inlet was developed with low dead-volume, allowing the use of a very low conductance limit interface. Gas flows through the ConFlat flange at a relatively high flow rate at orders of magnitude greater than through the conductance limit. The small flow goes through a conductance limit that is a double-sided ConFlat.

Naylor, Guy↗
NASA NTRS2011

The Evolution of Utilizing Manual Throttles to Avoid Excessively Low LH2 NPSP at the SSME Inlet

In the late 1970s, years before the Space Shuttle flew its maiden voyage, it was understood low liquid hydrogen (LH2) Net Positive Suction Pressure (NPSP) at the inlet to the Space Shuttle Main Engine (SSME) could have adverse effects on engine operation. A number of failures within both the External Tank (ET) and the Orbiter Main Propulsion System (MPS) could result in a low LH2 NPSP condition, which at extremely low levels can result in cavitation of SSME turbomachinery. Operational workarounds were developed to take advantage of the onboard crew s ability to manually throttle down the SSMEs (via the Pilot s Speedbrake/Throttle Controller), which alleviated the low LH2 NPSP condition. Manually throttling the SSME to a lower power level resulted in an increase in NPSP, mainly due to the reduction in frictional flow losses while at the lower throttle setting. Early in the Space Shuttle Program s history, the relevant Flight Rule for the Booster flight controllers in Mission Control did not distinguish between ET and Orbiter MPS failures and the same crew action was taken for both. However, after a review of all Booster operational techniques following the Challenger disaster in the late 1980s, it was determined manually throttling the SSME to a lower power was only effective for Orbiter MPS failures and the Flight Rule was updated to reflect this change. The Flight Rule and associated crew actions initially called for a single throttle step to minimum power level when a low threshold for NPSP was met. As engineers refined their understanding of the NPSP requirements for the SSME (through a robust testing program), the operational techniques evolved to take advantage of the additional capabilities. This paper will examine the evolution of the Flight rule and associated procedure and how increases in knowledge about the SSME and the Space Shuttle vehicle as a whole have helped shape their development. What once was a single throttle step when NPSP decreased to a certain threshold has now become three throttle steps, each occurring at a lower NPSP threshold. Additionally the procedure, which for early Space Shuttle missions required a Return-to-Launch-Site abort, now results in a nominal Main Engine Cut Off and no loss of mission objectives.

Henfling, Rick↗
NASA NTRS2011

Conical Probe Calibration and Wind Tunnel Data Analysis of the Channeled Centerbody Inlet Experiment

For a multi-hole test probe undergoing wind tunnel tests, the resulting data needs to be analyzed for any significant trends. These trends include relating the pressure distributions, the geometric orientation, and the local velocity vector to one another. However, experimental runs always involve some sort of error. As a result, a calibration procedure is required to compensate for this error. For this case, it is the misalignment bias angles resulting from the distortion associated with the angularity of the test probe or the local velocity vector. Through a series of calibration steps presented here, the angular biases are determined and removed from the data sets. By removing the misalignment, smoother pressure distributions contribute to more accurate experimental results, which in turn could be then compared to theoretical and actual in-flight results to derive any similarities. Error analyses will also be performed to verify the accuracy of the calibration error reduction. The resulting calibrated data will be implemented into an in-flight RTF script that will output critical flight parameters during future CCIE experimental test runs. All of these tasks are associated with and in contribution to NASA Dryden Flight Research Center s F-15B Research Testbed s Small Business Innovation Research of the Channeled Centerbody Inlet Experiment.

Truong, Samson Siu↗
NASA NTRS2012

Research on Supersonic Inlet Bleed

Phase I data results of the Fundamental Inlet Bleed Experiments project at NASA Glenn Research Center (GRC) are presented which include flow coefficient results for two single-hole boundary-layer bleed configurations. The bleed configurations tested are round holes at inclination angles of 90deg and 20deg both having length-to-diameter ratios of 2.0. Results were obtained at freestream Mach numbers of 1.33, 1.62, 1.98, 2.46, and 2.92 and unit Reynolds numbers of 0.984, 1.89, and 2.46 10(exp 7)/m. Approach boundary-layer data are presented for each flow condition and the flow coefficient results are compared to existing multi-hole data obtained under similar conditions. For the 90deg hole, the single and multi-hole distributions agree fairly well with the exception that under supercritical operation, the multi-hole data chokes at higher flow coefficient levels. This behavior is also observed for the 20deg hole but to a lesser extent. The 20deg hole also shows a markedly different characteristic at subcritical operation. Also presented are preliminary results of a Computational Fluid Dynamics (CFD) analysis of both configurations at the Mach 1.33 and a unit Reynolds number of 2.46 10(exp 7)/m. Comparison of the results shows the agreement to be very good.

Davis, David O.↗
NASA NTRS2013

The NTF Inlet Guide Vanes Thermal Gradient Problem and Its Mitigation

The National Transonic Facility (NTF) utilizes Inlet Guide Vanes (IGV) to provide precise, quick response Mach number control for the tunnel. During cryogenic operations, the massive IGV structure can experience large thermal gradients, measured as "Delta T or (Delta)T", between the IGV ring and its support structure called the transfer case. If these temperature gradients are too large, the IGV structure can be stressed beyond its safety limit and cease operation. In recent years, (Delta)T readings exceeding the prescribed safety limits were observed frequently during cryogenic operations, particularly during model access. The tactical operation methods of the tunnel to minimize (Delta)T did not always succeed. One obvious option to remedy this condition is to warm up the IGV structure by disabling the main drive operation, but this "natural" warm up method can takes days in some cases, resulting in productivity loss. This paper documents the thermal gradient problem associated with the IGV structure during cryogenic operation and how the facility has recently achieved an acceptable mitigation which has resulted in improved efficiency of operations.

Venkat, Venki S.↗
NASA NTRS2013

Aerodynamic Measurements of a Variable-Speed Power-Turbine Blade Section in a Transonic Turbine Cascade at Low Inlet Turbulence

Aerodynamic measurements obtained in a transonic linear cascade were used to assess the impact of large incidence angle and Reynolds number variations on the 3-D flow field and midspan loss and turning of a 2-D section of a variable-speed power-turbine (VSPT) rotor blade. Steady-state data were obtained for ten incidence angles ranging from +15.8 deg to −51.0 deg. At each angle, data were acquired at five flow conditions with the exit Reynolds number (based on axial chord) varying over an order-of-magnitude from 2.12×10(exp 5) to 2.12×10(exp 6). Data were obtained at the design exit Mach number of 0.72 and at a reduced exit Mach number of 0.35 as required to achieve the lowest Reynolds number. Midspan total-pressure and exit flow angle data were acquired using a five-hole pitch/yaw probe surveyed on a plane located 7.0 percent axial chord downstream of the blade trailing edge plane. The survey spanned three blade passages. Additionally, three-dimensional half-span flow fields were examined with additional probe survey data acquired at 26 span locations for two key incidence angles of +5.8 deg and −36.7 deg. Survey data near the endwall were acquired with a three-hole boundary-layer probe. The data were integrated to determine average exit total-pressure and flow angle as functions of incidence and flow conditions. The data set also includes blade static pressures measured on four spanwise planes and endwall static pressures. Tests were conducted in the NASA Glenn Transonic Turbine Blade Cascade Facility. The measurements reflect strong secondary flows associated with the high aerodynamic loading levels at large positive incidence angles and an increase in loss levels with decreasing Reynolds number. The secondary flows decrease with negative incidence as the blade becomes unloaded. Transitional flow is admitted in this low inlet turbulence dataset, making it a challenging CFD test case. The dataset will be used to advance understanding of the aerodynamic challenges associated with maintaining efficient power turbine operation over a wide shaft-speed range. deg

Flegel-McVetta, Ashlie B.↗
NASA NTRS2016

Experimental Evaluation of Inlet Distortion on an Ejector Powered Hybrid Wing Body at Take-off and Landing Conditions

As part of the NASA Environmentally Responsible Aircraft project, an ultra high bypass ratio engine integration on a hybrid wing body demonstration was planned. The goal was to include engine and airframe integration concepts that reduced fuel consumption by at least 50% while still reducing noise 42 db cumulative on the ground. Since the engines would be mounted on the upper surface of the aft body of the aircraft, the inlets may be susceptible to vortex ingestion from the wing leading edge at high angles of attack and sideslip, and separated wing/body flow. Consequently, experimental and computational studies were conducted to collect flow surveys useful for characterizing engine operability. The wind tunnel tests were conducted at two NASA facilities, the 14- by 22-foot at NASA Langley and the 40- by 80-foot at NASA Ames Research Center. The test results included in this paper show that the distortion and pressure recovery levels were acceptable for engine operability. The CFD studies conducted to compare to experimental data showed excellent agreement for the angle of attacks examined, although failed to match the low speed experimental data at high sideslip angles.

Carter, Melissa B.↗
NASA NTRS2017

Multipurpose Rotating Rake Arrays for Integrated Inlet and Fan Stage Performance Measurement

Low-pressure-ratio fan stage performance measurement requires precise measurement of conditions upstream and downstream of the fan stage. This presentation will discuss the rotating rake arrays used for the recent Boundary Layer Ingesting Inlet-Distortion-Tolerant Fan experiment in the NASA Glenn 8 by 6 foot wind tunnel. To achieve precise measurements, simulations of the rake sampling from pre-test CFD (Computerized Fluid Dynamics) solutions were used to optimize the number and locations of rake instruments.

performanc↗