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At least 379 records · Page 21
All sourcesNASA NTRSDOE OSTI
NASA NTRSDate not supplied

Inlet Vortex Generator Design for the X-59 Low Boom Flight Demonstrator

An array of vortex generators was designed for the X-59 Low Boom Flight Demonstrator aircraft inlet. Prior to the a High Speed Inlet DistortionTest, concern regarding potentially high measurements of steady state or dynamic distortion prompted an effort to examine the use of passive flow control to reduce distortion levels. Design of experiments methodology was used to develop two matrices of vortex generators placed inside the inlet serpentine duct and upstream of the inlet on the compression surface, varying on several geometric parameters. Reynolds- Averaged Navier-Stokes simulations were performed using the FUN3D flow solver at Mach 1.45 to analyze the impact of each configuration on total pressure recovery, circumferential distortion, and radial distortion. Results fromsimulations of the baseline aircraft are compared with data acquired in the NASA Glenn Research Center’s 8- by 6-foot SupersonicWind Tunnel. A vortex generator configuration with vanes placed upstream of the inlet terminal shock was shown to significantly increase total pressure recovery and decrease circumferential distortion with no penalty to radial distortion levels. Inlet mass flow sweeps at Mach 1.45, 0.6, and 0.3 indicated no detrimental effects from the vanes at off-design conditions.

Vortex generators↗
NASA NTRS1948

Ram-recovery Characteristics of NACA Submerged Inlets at High Subsonic Speeds

Results are presented of an experimental investigation of the characteristics of NACA submerged inlets on a model of a fighter airplane for Mach numbers from 0.30 to 0.875. The effects on the ram-recovery ratio at the inlets of Mach number, angle of attack, boundary-layer thickness on the fuselage, inlet location, and boundary-layer deflectors are shown. The data indicate only a slight decrease in ram-recovery ratio for the inlets ahead of or just behind the wing leading edge as Mach number increased, but showed large decreases at high Mach numbers for the inlets aft of the point of maximum thickness of the wing.

INLETS, SIDE - DUCTED BODIES↗
NASA NTRS1952

Performance of supersonic scoop inlets

A brief summary is made of the performance of a series of variable-geometry type scoop inlets investigated at the Lewis Laboratory on a model of the X-3 airplane fuselage. Data are presented for a range of inlet mass-flow ratios over a Mach number range from 0 to 2 and from 0 degrees to 12 degrees angle of attack. Rounded-lip inlets are found to give satisfactory performance to a Mach number of 1.5. Use of sharp rather than blunt lip configurations is shown to provide considerable gains in available thrust at Mach number 2.0 but penalize take-off performance to the extent that auxiliary air intakes may be required. No one inlet type was found to be superior over the entire range of operating variables. Thus, the choice of a specific inlet design will be influenced by structural and mechanical as well as aerodynamic considerations.

DIFFUSERS, SUPERSONIC↗
NASA NTRSDate not supplied

Validation of Noise Propagation for Lined Inlets

Commercial finite element modeling tools are validated for use on inlet radiated noise predictions for an acoustically lined subsonic aircraft inlet tested in the NASA Glenn 9-foot by 15-foot low-speed wind tunnel. An in-depth review and prediction of the acoustic liner properties on the Advanced Ducted Propulsor were conducted for use in the noise predictions. Agreement between the inlet radiated noise predictions to the measured tones at approach was used as the validation for the prediction tools. Similar follow-up predictions were conducted for an acoustically lined supersonic inlet. A follow-up investigation on the effect of the acoustic liner placement within the supersonic inlet was undertaken.

Acoustics↗
NASA NTRS2025

Influence of Actuator Movement Rate During Acquisition of Inlet Characterization Data

This report considers using a process to acquire performance data for supersonic mixed-compression inlets that is faster and less costly than the previous process. In the previous process for acquiring data during wind tunnel experiments, actuators for backpressuring an inlet moved in discrete increments, paused for dynamics to dampen, measured data for 5 s, recorded the data, and then reduced the data to graphic plots of pressure recovery versus mass capture ratio. This discrete data acquisition procedure was methodical, reliable, and accurate. However, acquiring data to characterize the inlet consumed significant wind-on time. Wind tunnel operation was costly when considering maintenance, energy, and human resources (operators and research team); therefore, the price for each data point was high. This report presents a quicker method to obtain all measurements, one which can also be used in blowdown wind tunnels. This approach, identified as the dynamic inlet characteristic data acquisition procedure, uses in situ dynamic high-speed pressure sensors to measure pressure signals. The backpressuring actuator moves continuously, and measurements from high-speed pressure sensors are recorded as the actuator progresses to and past the incremental points. Despite its advantages, the approach raised the following questions: (1) Will the new process be as accurate as the previous process? and (2) Are there any hidden costs? These concerns are addressed in this report. This report reviews a study that was undertaken to compare wind tunnel experiment data taken using the previous discrete data acquisition procedure with data recorded using the dynamic inlet characterization data acquisition procedure. In this report, the wind tunnel experiment and the study objectives are described. The processes to reduce the data for analysis are explained, and a discussion of the results is given. Finally, conclusions from this study that may guide further work in this area are presented.

Inlet Characterization Curves↗
NASA NTRS1953

Force and pressure-recovery characteristics at supersonic speeds of a conical nose inlet with bypasses discharging outward from the body axis

Aerodynamic and performance characteristics of a conical spike nacelle-type inlet with two bypasses are presented at Mach numbers of 1.6, 1.8, and 2.0 for angles of attach up to 90 degrees. The bypasses were located 6 inlet diameters downstream of the inlet and were designed to discharge the bypass mass flow outward from the body axis. The inlet was designed to attain a mass-flow ratio of unity at a Mach number of 2.0. It is shown that discharging the bypass mass flow outward from the body nearly doubles the critical drag of a similar configuration but with bypass discharge in an axial direction. As a result of this greater drag, the net force on the model in the flight direction is reduced when comparison is made with the axial discharge case. The lift and pitching-moment coefficients are slightly higher than those for a configuration without bypasses. Approximately 25 % of the maximum inlet mass flow was discharged through the bypasses, and the pressure-recovery and mass-flow characteristics were in qualitative and quantitative agreement with the results of an investigation of a similar configuration with axial discharge.

DIFFUSERS, SUBSONIC↗
NASA NTRS2015

Inward-Turning Streamline-Traced Inlet Design Method for Low-Boom, Low-Drag Applications

A new design method for inward-turning, streamline-traced inlets is presented. Resulting designs are intended for moderate supersonic, low-drag, low-boom applications such as that required for NASA's proposed low-boom flight demonstration aircraft. A critical feature of these designs is the internal cowl lip angle that allows for little or no flow turning on the outer nacelle. Present methods using conical-flow Busemann parent flowfields have simply truncated, or otherwise modified the stream-traced contours to include this internal cowl angle. Such modifications disrupt the parent flowfield, reducing inlet performance and flow uniformity. The method presented herein merges a conical flowfield that includes a leading shock with a truncated Busemann flowfield in a manner that minimizes unwanted interactions. A leading internal cowl angle is now inherent in the parent flowfield, and inlet contours traced from this flowfield retain its high performance and good flow uniformity. CFD analysis of a candidate inlet design is presented that verifies the design technique, and reveals a starting issue with the basic geometry. A minor modification to the cowl lip region is shown to eliminate this phenomenon, thereby allowing starting and smooth transition to sub-critical operation as back-pressure is increased. An inlet critical-point total pressure recovery of 96 is achieved based on CFD results for a Mach 1.7 freestream design. Correction for boundary-layer displacement thickness, and sizing for a given engine airflow requirement are also discussed.

low sonic boom↗
NASA NTRS2015

Enhanced Performance of Streamline-Traced External-Compression Supersonic Inlets

A computational design study was conducted to enhance the aerodynamic performance of streamline-traced, external-compression inlets for Mach 1.6. Compared to traditional external-compression, two-dimensional and axisymmetric inlets, streamline-traced inlets promise reduced cowl wave drag and sonic boom, but at the expense of reduced total pressure recovery and increased total pressure distortion. The current study explored a new parent flowfield for the streamline tracing and several variations of inlet design factors, including the axial displacement and angle of the subsonic cowl lip, the vertical placement of the engine axis, and the use of porous bleed in the subsonic diffuser. The performance was enhanced over that of an earlier streamline-traced inlet such as to increase the total pressure recovery and reduce total pressure distortion.

performance↗
NASA NTRS2016

Vortex Generators in a Two-Dimensional, External-Compression Supersonic Inlet

Vortex generators within a two-dimensional, external-compression supersonic inlet for Mach 1.6 were investigated to determine their ability to increase total pressure recovery, reduce total pressure distortion, and improve the boundary layer. The vortex generators studied included vanes and ramps. The geometric factors of the vortex generators studied included height, length, spacing, and positions upstream and downstream of the inlet terminal shock. The flow through the inlet was simulated through the computational solution of the steady-state Reynolds-averaged Navier-Stokes equations on multi-block, structured grids. The vortex generators were simulated by either gridding the geometry of the vortex generators or modeling the vortices generated by the vortex generators. The inlet performance was characterized by the inlet total pressure recovery, total pressure distortion, and incompressible shape factor of the boundary-layer at the engine face. The results suggested that downstream vanes reduced the distortion and improved the boundary layer. The height of the vortex generators had the greatest effect of the geometric factors.

computational fluid dynamics↗
NASA NTRS2017

Vortex Generators in a Streamline-Traced, External-Compression Supersonic Inlet

Vortex generators within a streamline-traced, external-compression supersonic inlet for Mach 1.66 were investigated to determine their ability to increase total pressure recovery and reduce total pressure distortion. The vortex generators studied were rectangular vanes arranged in counter-rotating and co-rotating arrays. The vane geometric factors of interest included height, length, spacing, angle-of-incidence, and positions upstream and downstream of the inlet terminal shock. The flow through the inlet was simulated numerically through the solution of the steady-state, Reynolds-averaged Navier-Stokes equations on multi-block, structured grids using the Wind-US flow solver. The vanes were simulated using a vortex generator model. The inlet performance was characterized by the inlet total pressure recovery and the radial and circumferential total pressure distortion indices at the engine face. Design of experiments and statistical analysis methods were applied to quantify the effect of the geometric factors of the vanes and search for optimal vane arrays. Co-rotating vane arrays with negative angles-of-incidence positioned on the supersonic diffuser were effective in sweeping low-momentum flow from the top toward the sides of the subsonic diffuser. This distributed the low-momentum flow more evenly about the circumference of the subsonic diffuser and reduced distortion. Co-rotating vane arrays with negative angles-of-incidence or counter-rotating vane arrays positioned downstream of the terminal shock were effective in mixing higher-momentum flow with lower-momentum flow to increase recovery and decrease distortion. A strategy of combining a co-rotating vane array on the supersonic diffuser with a counter-rotating vane array on the subsonic diffuser was effective in increasing recovery and reducing distortion.

computational fluid dynamics↗
NASA NTRSDate not supplied

Inlet Performance of the NFAC 1/50th-scale 80- by 120- Foot Wind Tunnel

The National Full-Scale Aerodynamics Complex 80- by 120-Foot Wind Tunnel (80x120) was dedicated in 1987 and rated at 100 knots for full-scale aircraft testing at NASA Ames Research Center. The 80x120 is the world’s largest wind tunnel, designed as an open circuit tunnel with a large aerodynamically treated inlet open to the ambient atmospheric air. In 2017, damage was sustained within the wind tunnel drive system, opening a window to do testing using the existing 1/50th-scale model of the 80- by 120-Foot Wind Tunnel within the full-scale 80- by 120-Foot Wind tunnel test section. The objective of the research was to quantify the turbulence levels within the 1/50th-scale test section as a function of onset atmospheric wind direction (± 90 deg from tunnel center-line), variable test section speed (5 – 50 m/s) and purposeful obtrusion of wind flow into the inlet. The model wind tunnel inlet, contraction, and test section are geometrically identical to that of the full-scale wind tunnel and model testing provides aerodynamic performance characteristics under controlled test conditions allowing for insight into the full-scale test section flow quality. The test section turbulence levels are minimally affected by the onset direction of the ambient atmospheric wind, but are dramatically affected by the speed in the tunnel while operated in the presence of winds. Original design specifications were axial/vertical/lateral turbulence ≤ 0.5% at maximum test section speed, though early full scale tunnel testing determined that lateral turbulence would be ≤ 0.6%. For test section speeds ≥ 30 m/s the tunnel is within the design specification limits. Between 5 m/s and 30 m/s, the test section turbulence levels are dependent on the onset wind direction and test section speed where test section turbulence in the axial, vertical and lateral directions was seen to be between 0.5% and 1% and, at times, greater than 1%. Finally, testing was performed with blockage designs at the inlet to disrupt the wind flow quality entering the tunnel contraction zone in an attempt to create higher levels of turbulence for high turbulent test conditions simulating the earth’s boundary layer. The highest turbulence levels measured were 6% in the axial direction by use of large spires designed to obstruct ≈ 50% of the inlet area.

Inlet Performance↗
NASA NTRS2021

Experimental Results and Interfacial Lift-off Model Predictions of Critical Heat Flux for Flow Boiling with Subcooled Inlet Conditions – In Preparation for Experiments Onboard the International Space Station

This study investigates critical heat flux (CHF) for subcooled flow boiling of n-Perfluorohexane based on results of pre-launch Earth-gravity Mission Sequence Tests (MSTs) of the Flow Boiling and Condensation Experiment (FBCE), which was launched to the International Space Station (ISS) in August 2021. CHF measurements were made in a rectangular channel having a 2.5 mm by 5 mm cross-section and a 114.6-mm long heated segment. Both single-sided and double-sided heating were tested in vertical upflow in Earth gravity for a variety of inlet conditions. The inlet subcooling was varied in the range of 0.4 – 32.0°C and encompassed both near-saturated and highly subcooled conditions. Experimental trends and high-speed video records were investigated to better understand the mechanism of CHF. Overall trends show CHF increases as flow rate and/or inlet subcooling are increased. Flow features from the events around CHF justify the applicability of the Interfacial Lift-off Model and the determination of limiting criteria for its application. The present experimental data are combined with prior databases for various flow orientations with respect to Earth gravity and microgravity data collected on parabolic flights. Predictions are made using the Interfacial Lift-off Model for this consolidated subcooled-inlet FBCE-CHF database. A heat utility ratio was included in the model to capture the effects of subcooling and corresponding thermodynamic non-equilibrium. An overall mean absolute error of 19.04% indicates good predictive capability of the model for both heating configurations, different gravity environments, and a wide range of inlet subcooling.

flow boiling↗
NASA NTRS2022

Flow Visualization, Heat Transfer, and Critical Heat Flux of Flow Boiling in Earth Gravity with Saturated Liquid-Vapor Mixture Inlet Conditions – In Preparation for Experiments Onboard the International Space Station

This study investigates flow boiling of n-Perfluorohexane with saturated two-phase mixture inlet in a rectangular channel of dimensions 114.6-mm heated length, 2.5-mm width, and 5 mm-height. The experiments were performed as part of the Mission Sequence Testing of the Flow Boiling and Condensation Experiment’s (FBCE) Flow Boiling Module (FBM) in the vertical upflow configuration in Earth gravity using the same experimental system that was launched to the International Space Station (ISS) in August 2021. The operating parameters varied are heating configuration (single- and double-sided), mass velocity (380 – 2400 kg/m2s), inlet quality (0.011 – 0.519), and inlet pressure (120 – 179 kPa). High-speed video photographs are presented to explain the two-phase flow patterns within the channel’s heated length. Flow patterns are constituted by low-density and high-density fronts moving along the channel, with the high-density fronts gradually reducing in length due to evaporation. Heat transfer results in terms of flow boiling curves, streamwise wall temperature profiles, streamwise heat transfer coefficient profiles, and average heat transfer coefficients are presented and trends discussed. CHF data from the present experiments are combined with prior databases to compile a consolidated FBCE-CHF database for saturated inlet to expand the ranges of operating conditions and include other flow orientations in Earth gravity. Experimental CHF trends are also discussed. The interfacial lift-off model shows a good CHF predictive accuracy evidenced by a mean absolute error of 11.97% for this consolidated database after constraining it to mass velocities greater than or equal to 500 kg/m2s. Finally, this study confirmed reliability of the upcoming ISS experiments for saturated inlet conditions and the collected Earth-gravity data will be compared to ISS microgravity data.

Flow boiling↗
NASA NTRSDate not supplied

FUN3D Analyses of the 5th AIAA Propulsion Aerodynamics Workshop Inlet Test Case

This work presents the results of FUN3D analyses that were performed for the 5th AIAA Propulsion Aerodynamics Workshop. The workshop was separated into two sections that focused on inlet and nozzle flows. This paper focuses on the inlet section of the workshop, which considered an axisymmetric inlet in close-proximity to the ground and subjected to a crosswind. Four levels of grid refinement were provided by the workshop committee. Participants were asked to perform simulations for three crosswind velocities on a minimum of three levels of grid refinement, for a total of nine required cases. This work employed an updated set of grids provided by the committee after completion of the workshop based on feedback from participants. The updated grids featured a larger computational domain and increased refinement near the ground vortex region. This work employed the three coarsest grid levels from the updated grid set. Note that an additional extension downstream of the inlet was added to the provided grids for this work to mitigate observed convergence issues. Additionally, a limited turbulence model study was performed for the highest value of crosswind velocity and the finest grid. The results show that grid convergence was only achieved for the case with the lowest value of crosswind velocity. However, the results were generally observed to improve, relative to the experimental data, with increasing grid refinement. The FUN3D predictions were observed to agree qualitatively with the provided experimental data. However, FUN3D exhibited a tendency to underpredict the minimum value of the radially averaged total pressure at the Aerodynamic Interface Plane, which is consistent with the general findings from the workshop. Overall, FUN3D was found to perform well relative to its peers for the PAW5 workshop inlet test case.

PAI↗
NASA NTRSDate not supplied

FUN3D Analyses of the 5th AIAA Propulsion Aerodynamics Workshop Inlet Test Case

This work presents the results of FUN3D analyses that were performed for the 5th AIAA Propulsion Aerodynamics Workshop. The workshop was separated into two sections that focused on inlet and nozzle flows. This paper focuses on the inlet section of the workshop, which considered an axisymmetric inlet in close-proximity to the ground and subjected to a crosswind. Four levels of grid refinement were provided by the workshop committee. Participants were asked to perform simulations for three crosswind velocities on a minimum of three levels of grid refinement, for a total of nine required cases. This work employed an updated set of grids provided by the committee after completion of the workshop based on feedback from participants. The updated grids featured a larger computational domain and increased refinement near the ground vortex region. This work employed the three coarsest grid levels from the updated grid set. Note that an additional extension downstream of the inlet was added to the provided grids for this work to mitigate observed convergence issues. Additionally, a limited turbulence model study was performed for the highest value of crosswind velocity and the finest grid. The results show that grid convergence was only achieved for the case with the lowest value of crosswind velocity. However, the results were generally observed to improve, relative to the experimental data, with increasing grid refinement. The FUN3D predictions were observed to agree qualitatively with the provided experimental data. However, FUN3D exhibited a tendency to underpredict the minimum value of the radially averaged total pressure at the Aerodynamic Interface Plane, which is consistent with the general findings from the workshop. Overall, FUN3D was found to perform well relative to its peers for the PAW5 workshop inlet test case.

PAI↗
NASA NTRS1952

Performance Characteristics of a Normal-shock Side Inlet Located Downstream of a Canard Control Surface at Mach Numbers of 1.5 and 1.8/

The performance characteristics of a downward canted normal-shock side (scoop) inlet located downstream of a triangular control surface are presented for free-stream Mach numbers of 1.5 and 1.8 in terms of total pressure recovery and mass flow ratio for various boundary-layer removal systems,angles of attack, control surface deflections and adverse yaw. An engine operating condition for a hypothetical turbojet engine is established, and the match point characteristics of the engine-inlet configuration are summarized. 520::It is shown that the diffuser performance increases with increased boundary-layer removal and decreases because of the presence of the wake from the forward control surface. At the higher angles of attack the wake passes over the inlet and does not affect the inlet performance. Adverse yaw reduces the total pressure recovery values below those for the unawed case. Magnitudes of the total pressure recovery were below the theoretical normal-shock recovery for the respective test Mach numbers.

INLETS, SIDE - DUCTED BODIES↗
DOE OSTI2021

Elevated Inlet Temperature Effects on the Operation of a Rotating Detonation Combustor

This experimental work explores the impact of elevated inlet temperatures on a H2/air operated rotating detonation combustor (RDC) without a nozzle. In practical systems, the temperature of the inlet air would be higher than that found in many laboratory settings. Inlet air pre-heating up to 480 K was accomplished by using an integrated pebble-bed heater. The operation of the RDC was found to switch from a single detonation wave at ambient temperature to two, counter-rotating detonation waves at higher temperatures for some of the conditions tested. In addition, the detonation wave properties (i.e., speed and pressure ratio) were impacted by the elevated inlet temperature. The cause of the observed changes in the operation dynamics was explored by evaluating the existence of secondary waves prior to the operational mode change and the changes in the cycle-averaged inlet Mach number.

Feleo, Alexander↗
NASA NTRS1957

Interference Effects of Fuselage-Stored Missiles on Inlet Duct Model of an Interceptor-Type Aircraft at Mach Numbers 1.5 to 1.9

The effect of missile armament on the performance of an interceptor-type aircraft model has been determined at Mach numbers 1.5, 1.7, and 1.9 and at angles of attack to 19 deg. With this configuration missiles were carried in a bay located on the bottom of the aircraft fuselage and mounted to a rotatable missile door. Rotation of the door then brought the missiles into the external or firing position. The aircraft model was characterized by triangular-shaped normal-shock inlets located at the wing roots. Relatively short and curved subsonic diffusers fed simulated twin side-by-side turbojet engines. Inasmuch as the missile bay extended considerably ahead of the inlet station, rotation of the missile door created considerable disturbance of the flow entering the inlets. In comparison with the internal missile arrangement, the external missile configurations increased the model lift, drag, and pitching moment. While the diffuser-exit flow distortion and static-pressure fluctuations were not greatly affected, diffuser total-pressure recovery was reduced as much as 0.058 at Mach number 1.9 for one missile configuration. The most detrimental effect of missile-door rotation occurred at the transient door positions, or with the door halfwzy between the missiles-in and -out conditions. At this door position the flow into the inlets was highly asymmetrical. Although the performance of both left and right ducts was generally reduced, the inlet duct on the cavity side of the missile door was most severely penalized, becoming unstable recovery losses and increases in flow distortion. The installation of fuselage fences along the missile bay was only partially effective in reducing these losses.

Piercy, T. G.↗