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Search indexed NASA NTRS and DOE OSTI research on propulsion, heat transfer, battery materials and energy systems. Follow report and document links to the original sources.

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At least 271 records · Page 15
All sourcesNASA NTRSDOE OSTI
NASA NTRS1984

Flow rates and pressure profiles for one to four axially alined Borda inlets

Choked flow rate and pressure profile data were taken on sequential, axially alined inlets of the Borda type. The configurations consisted of two to four inlets spaced 0.8 and 30 diameters apart. At a spacing of 30 diameters the reduced flow rate appeared to follow the simple empirical relation G sub r/G sub r,1=N(-b), where G sub r,1 is the reduced flow rate for a single inlet; N is the number of inlets; and b, which is weakly temperature dependent, is approximately 0.4. The relation is in reasonable agreement with an analysis of the N-inlet configuration. At a spacing of 30 diameters the pressure profiles dropped sharply at the entrance and partially recovered within each inlet somewhat independently of N. Jetting through the last Borda was common at low temperatures. At a spacing of 0.8 diameter fluid jetting was prevalent at low temperatures for each configuration studied and flow rates were the same as that for a single inlet.

Hendricks, R. C.↗
NASA NTRS1986

Study of hypersonic inlet flow fields with a three-dimensional Navier-Stokes code

A three-dimensional Navier-Stokes code is used to study the flow through an inlet with a rectangular capture area, but with circular or essentially circular internal cross-sections. The inlet may have a center body extending partially or fully over the length of the inlet. Calculations are made under a variety of geometrical constraints imposed on the inlet cross-sections to study the flow quality in the inlet. It is observed that for this class of inlets, the boundary layer has a tendency to thicken and separate readily due to the adverse pressure gradient caused by the continuous compression in the inlet. Since no experimental results are yet available, this paper has presented only the numerical results from the present analysis.

Kumar, A.↗
NASA NTRS1987

Electromagnetic fields backscattered from an s-shaped inlet cavity with an absorber coating on its inner walls

The EM backscatter from a two-dimensional S-shaped inlet cavity is analyzed using three different techniques, namely a hybrid combination of asymptotic high frequency and modal methods, an integral equation method, and the geometrical optics ray method, respectively. This inlet has a thin absorber coating on its perfectly conducting inner walls and the planar interior termination is made perfectly conducting. The effect of the absorber on the inner wall is treated via a perturbation scheme in the hybrid approach where it is assumed that the loss is sufficiently small for the method to be valid. The results are compared with the backscatter from a straight inlet cavity to evaluate the effect of offsetting the termination in the S-bend configuration such that it is not visible from the open end of the inlet. The envelope of the backscatter pattern for the straight inlet is always seen to peak around the forward axis due to the large return from the directly visible termination, and the pattern envelope tapers off away from the forward axis. Offsetting the termination causes the envelope of the backscatter pattern to flatten out, thereby reducing the return near the forward axis by several dB. The absorber coating reduces the pattern level of the straight inlet in directions away from the forward axis but has little effect on the peak near the axis; furthermore, the absorber coating is seen to consistently reduce the backscatter from the S-bend inlet for almost all incidence angles. The hybrid method gives excellent agreement with experimental data and with the integral equation solution, whereas, the geometrical optics ray tracing method is able to generally predict the average of the bachscatter pattern but not the pattern details.

Burkholder, R. J.↗
NASA NTRS1989

Analysis of the SSME HPOTP bearing inlet cavity

Analysis of the flow in the Space Shuttle Main Engine (SSME) high pressure oxygen turbopump (HPOTP) bearing no. 1 inlet cavity was completed in support of return-to-flight. With the incorporation of several design changes in the Phase 2 turbopump, rotordynamic stability of the pumps was enhanced, but the durability and life of the LOX-cooled bearings has decreased. During the post-Challenger SSME recertification, the causes of limited bearing durability were investigated. One topic addressed was the flow environment upstream of the pump-end bearing and the effect of seal exit swirl and a cavity anti-vortex rib on the bearing environment and life. The objective is to define the hydrodynamic environment upstream of the pump-end bearing and determine the effect of seal exit swirl and the anti-vortex rib on bearing inlet swirl. The problem was posed as an axisymmetric cavity flow with the computational domain extending from the seal exit to the bearing inlet. This domain was discretized with 22800 grid points. Boundary conditions were obtained from a 1-D model of the SSME coolant path. The inlet Mach number was 0.19 and the problem was solved with the CMINT code utilizing the Briley-McDonald/Beam-Warming algorithm with preconditioning to speed convergence at low Mach numbers. Three parametric cases with inlet swirl of 50 percent shaft speed (labyrinth seal), 20 percent shaft speed (damping seal), and no inlet swirl were considered. Computational results indicate large vortical flow structures in the cavity, with the labyrinth, damping, and no-swirl cases yielding bearing inlet swirl rates of 14, 10, and 9 percent of shaft speed, respectively. When these results were used as input to the SHABRETH bearing model, limited durability could not be explained by these small differences in swirl. Also, based on these results, a proposed design change for the cavity anti-vortex rib was not implemented by the SSME chief engineer.

Mcconnaughey, P. K.↗
NASA NTRS1992

Coupled numerical simulation of the external and engine inlet flows for the F-18 at large incidence

This paper presents a numerical simulation of the external and engine inlet flows for the F-18 aircraft at typical high-angle-of-attack flight conditions. Two engine inlet mass flow rates, corresponding to flight idle and maximum power, were computed. This was accomplished using a structured, overset grid technique to couple the external and internal grid systems. Reynolds-averaged Navier-Stokes solutions were obtained using an implicit, finite-differencing scheme. Results show a strong coupling of the external and engine inlet flows, especially at the maximum power setting. Increasing the mass flow rate through the inlet caused the primary vortex breakdown location to move downstream. This trend is also observed in flight tests performed on the F-18. A reversed flow region upstream of the inlet duct is visible in the faired-inlet and flight-idle computations. This flow reversal is not present in the maximum power setting computation. These large-scale changes in flow structure highlight the importance of simulating inlet conditions in high-angle-of-attack aircraft computations.

Murman, Scott M.↗
NASA NTRS1993

Analytical and experimental studies of a short compact subsonic diffuser for a two-dimensional supersonic inlet

An experimental study of a two-dimensional supersonic inlet with a short compact subsonic diffuser, length to exit diameter (dl/d) ratio of 1.25, was conducted to investigate the impact of the short diffuser on inlet performance at low speeds and to assess the diffuser subsonic performance for a simulated diffuser flow corresponding to high-speed inlet conditions near the design flight Mach number of 2.2. For the low-speed testing, a drooped lip was employed to improve the inlet performance at a high angle of attack. For the simulated high-speed testing, air was blown through slots or discrete nozzles as an active boundary-layer control. The results from the low-speed performance test were compared with the results from a previous test program on the same inlet with a long subsonic diffuser (dl/d = 4.5). The comparison indicates that inlet recovery was not affected by the use of the short diffuser for either the baseline (no droop) or the drooped cowl lip configuration. However, the inlet baseline distortion for the short diffuser configuration was substantially higher than for the long diffuser. A comparison of the two configurations with a 70 deg drooped lip showed no significant difference in distortion. For the portion of the experimental program in which diffuser conditions for high-speed flight were simulated, diffuser-induced flow separation occurred. This separation was predicted from an analytical study that used the Hess potential flow panel method and the Herring two-dimensional boundary-layer analysis computer codes. The flow separated mainly on the diffuser ramp. Subsequent tests in which boundary-control systems were utilized showed that blowing with either slots or discrete nozzles could suppress the flow separation in the short subsonic diffuser, thereby substantially improving the diffuser performance.

Iek, Chanthy↗
NASA NTRS1993

The 3-D viscous flow CFD analysis of the propeller effect on an advanced ducted propeller subsonic inlet

A time marching Navier-Stokes code called PARC3D was used to study the 3-D viscous flow associated with an advanced ducted propeller (ADP) subsonic inlet at take-off operating conditions. At a free stream Mach number of 0.2, experimental data for the inlet-with-propeller test model indicated that the airflow was attached on the cowl windward lip at an angle of attack of 25 degrees became unstable at 29 degrees, and separated at 30 degrees. An experimental study with a similar inlet and with no propeller (through-flow) indicated that flow separation occurred at an angle of attack a few degrees below the value observed when the inlet was tested with the propeller. This tends to indicate that the propeller exerts a favorable effect on the inlet performance. During the through-flow experiment a stationary blockage device was used to successfully simulate the propeller effect on the inlet flow field at angles of attack. In the present numerical study, this flow blockage was modeled via a PARC3D computational boundary condition (BC) called the screen BC. The principle formulation of this BC was based on the one-and-half dimension actuator disk theory. This screen BC was applied at the inlet propeller face station of the computational grid. Numerical results were obtained with and without the screen BC. The application of the screen BC in this numerical study provided results which are similar to the results of past experimental efforts in which either the blockage device or the propeller was used.

Iek, Chanthy↗
NASA NTRS1948

The Development and Application of High-Critical-Speed Nose Inlets

An analysis of the nose-inlet shapes developed in previous investigations to represent the optimum from the standpoint of critical speed has shown that marked similarity exists between the nondimensional profiles of inlets which have widely different proportions and critical speeds. With the nondimensional similarity of such profiles established, the large differences in the critical speeds of these nose inlets must be a function of their proportions. An investigation was undertaken in the Langley 8-foot high-speed tunnel to establish the effects of nose-inlet proportions on critical Mach number to develop a rational method for the design of high-critical-speed nose inlets to meet desired requirements. The test results data have been arranged in the form of design charts from NACA 1-series nose-inlet proportions and can be selected for given values of critical Mach number and airflow quantity. Examples of nose-inlet selections are presented for a typical jet-propulsion installation (critical Mach number of 0.83) and for two conventional radial-engine installations (critical Mach number of 0.76).

Baals, Donald D↗
NASA NTRS1950

Two-dimensional symmetrical inlets with external compression

The purpose of inlets like, for instance, those of air-cooled radiators and scoops is to take a certain air quantity out of the free stream and to partly convert the free-stream velocity into pressure. In the extreme case this pressure conversion may occur either entirely in the interior of the inlet (inlet with internal compression) or entirely in the free stream ahead of the inlet (inlet with external compression). In this report a theory for two-dimensional inlets with external compression is developed and illustrated by numerical examples. Intermediary forms between inlets with internal and external compression which can be derived from the latter are briefly discussed.

FLOW - DIFFUSERS↗
NASA NTRS1994

Installed F/A-18 inlet flow calculations at 30 degrees angle-of-attack: A comparative study

NASA Lewis is currently engaged in a research effort as a team member of the High Alpha Technology Program (HATP) within NASA. This program utilizes a specially equipped F/A-18, the High Alpha Research Vehicle (HARV), in an ambitious effort to improve the maneuverability of high-performance military aircraft at low subsonic speed, high angle of attack conditions. The overall objective of the Lewis effort is to develop inlet technology that will ensure efficient airflow delivery to the engine during these maneuvers. One part of the Lewis approach utilizes computational fluid dynamics codes to predict the installed performance of inlets for these highly maneuverable aircraft. Full Navier-Stokes (FNS) calculations on the installed F/A-18 inlet at 30 degrees angle of attack, 0 degrees yaw, and a freestream Mach number of 0.2 have been obtained in this study using an algebraic turbulence model with two grids (original and revised). Results obtained with the original grid were used to determine where further grid refinements and additional geometry were needed. In order to account properly for the external effects, the forebody, leading edge extension (LEX), ramp, and wing were included with inlet geometry. In the original grid, the diverter, LEX slot, and leading edge flap were not included due to insufficient geometry definition, but were included in a revised grid. In addition, a thin-layer Navier-Stokes (TLNS) code is used with the revised grid and the numerical results are compared to those obtained with the FNS code. The TLNS code was used to evaluate the effects on the solution using a code with more recent CFD developments such as upwinding with TVD schemes versus central differencing with artificial dissipation. The calculations are compared to a limited amount of available experimental data. The predicted forebody/fuselage surface static pressures compared well with data of all solutions. The predicted trajectory of the vortex generated under the LEX was different for each solution. These discrepancies are attributed to differences in the grid resolution and turbulence modeling. All solutions predict that this vortex is ingested by the inlet. The predicted inlet total pressure recoveries are lower than data and the distortions are higher than data. The results obtained with the revised grid were significantly improved from the original grid results. The original grid results indicated the ingested vortex migrated to the engine face and caused additional distortions to those already present due to secondary flow development. The revised grid results indicate that the ingested vortex is dissipated along the inlet duct inboard wall. The TLNS results indicate the flow at the engine face was much more distorted than the FNS results and is attributed to the pole boundary condition introducing numerical distortions into the flow field.

Smith, C. Frederic↗
NASA NTRS1995

PARC3D calculations of the F/A-18A HARV inlet vortex generators

NASA Lewis Research Center is currently engaged in a research effort as a team member of the High Alpha Technology Program within the NASA agency. This program uses a specially-equipped F/A-18A aircraft called the High Alpha Research Vehicle (HARV), in an effort to improve the maneuverability of high performance military aircraft at low-subsonic-speed, high-angle-of-attack conditions. The overall objective of the NASA Lewis effort is to develop inlet analysis technology towards efficient airflow delivery to the engine during these maneuvers. One portion of this inlet analysis technology uses computational fluid dynamics to predict installed inlet performance. Most of the F/A-18A HARV geometry, which includes the ramp/splitter plate, side diverter and slot, inlet lip and upper diverter, and deflected leading-edge flap has been modeled. The empennage and rear fuselage have not. A pair of vortex generators located on the bottom wall of the inlet were not modeled initially. These vortex generators were installed to alleviate any flow separation that may be induced by the wheel well protrusion into the inlet wall. Calculations completed with the PARC3D code showed that the pressure recovery has been underpredicted and the flow distortion over-predicted. To improve the correlation of PARC3D predictions with flight and wind tunnel tests, the vortex generators were included in the grid geometry and the results are presented in this report. The grid totals 27 blocks or 1.3 million grid points for the half model, which includes the vortex generator grid blocks. Two flight cases were calculated, a high speed case with a Mach number of 0.8 and angle of attack of 3.4; and a low speed case with a Mach number of 0.43 and angle of attack of 32.2. The vortex generators have a significant effect on the inlet boundary layers at high speed, low angle of attack; and have no effect at low speed, high angle of attack.

Podleski, Steve D.↗
NASA NTRS1997

Inlet Distortion for an F/A-18A Aircraft During Steady Aerodynamic Conditions up to 60 deg Angle of Attack

The effects of high-angle-of-attack flight on aircraft inlet aerodynamic characteristics were investigated at NASA Dryden Flight Research Center, Edwards, California, as part of NASA's High Alpha Technology Program. The highly instrumented F/A-18A High Alpha Research Vehicle was used for this research. A newly designed inlet total-pressure rake was installed in front of the starboard F404-GE-400 engine to measure inlet recovery and distortion characteristics. One objective was to determine inlet total-pressure characteristics at steady high-angle-of-attack conditions. Other objectives include assessing whether significant differences exist in inlet distortion between rapid angle-of-attack maneuvers and corresponding steady aerodynamic conditions, assessing inlet characteristics during aircraft departures, providing data for developing and verifying computational fluid dynamic codes, and calculating engine airflow using five methods. This paper addresses the first objective by summarizing results of 79 flight maneuvers at steady aerodynamic conditions, ranging from -10 deg to 60 deg angle of attack and from -8 deg to 11 deg angle of sideslip at Mach 0.3 and 0.4. These data and the associated database have been rigorously validated to establish a foundation for understanding inlet characteristics at high angle of attack.

Walsh, Kevin R.↗
NASA NTRS1998

Summary of Inlet Characteristics of the F/A-18A High Alpha Research Vehicle

Effects of high-angle-of-attack flight on aircraft inlet aerodynamic characteristics were investigated at NASA Dryden Flight Research Center as part of NASA's High Alpha Technology Program. The highly instrumented F/A-18A High Alpha Research Vehicle was used for this research. A newly designed inlet total-pressure rake was installed in front of the right-hand F404-GE-400 engine to measure inlet recovery and distortion characteristics. Objectives included: (1) determining the inlet total-pressure characteristics at steady high-angle-of-attack conditions; (2) assessing if inlet distortion is significantly different between rapid angle-of-attack maneuvers and corresponding steady aerodynamic conditions; (3) assessing inlet characteristics during aircraft departures; (4) providing data for developing and verifying computational fluid dynamic codes; and (5) calculating engine airflow using four methods for comparison with a reference method. This paper describes the results obtained from this investigation. These data and the associated database were rigorously validated to establish the foundation for understanding inlet characteristics at high angle of attack.

Walsh, Kevin↗
NASA NTRS2002

Control of Inflow Distortion in a Scarf Inlet

The scarf inlet has the potential to reduce aircraft inlet noise radiation to the ground by reflecting it into the space above the engine. Without forward motion of the engine, the non-symmetry of the inlet causes inflow distortion which generates noise that is greater than the noise reduction of the scarf. However, acoustic evaluations of aircraft engines are often done on static test stands. A method to reduce inflow distortion by boundary layer suction is proposed and evaluated using a model of a high bypass ratio engine located in an anechoic chamber. The design goal of the flow control system is to make the inflow to the inlet circumferentially uniform and to eliminate reversed flow. This minimizes the inflow distortion and allows for acoustic evaluation of the scarf inlet on a static test stand. The inlet boundary layer suction effectiveness is evaluated both by aerodynamic and by acoustic measurements. Although the design goal is not met, the control system is found to have a beneficial effect on the engine operation, reducing blade stall and speed variation. This is quantified by two acoustic benefits, reduction both of the variability of tone noise and of the low frequency wideband noise due to the inflow distortion. It is felt that a compromise in the manufacture of the control hardware contributes to the inability of the control system to perform as expected from the analysis. The control system with sufficient authority is felt to have the potential to permit reliable acoustic testing in a static configuration of engines with non-symmetric inlets. Because the control system can improve operation of the engine, it may also have the potential to reduce noise and vibration and enhance engine longevity during low speed ground operations in the terminal area.

Gerhold, Carl H.↗
NASA NTRS2004

Active Flow Control on a Boundary-Layer-Ingesting Inlet

Boundary layer ingestion (BLI) is explored as means to improve overall system performance for Blended Wing Body configuration. The benefits of BLI for vehicle system performance benefit are assessed with a process derived from first principles suitable for highly-integrated propulsion systems. This performance evaluation process provides framework within which to assess the benefits of an integrated BLI inlet and lays the groundwork for higher-fidelity systems studies. The results of the system study show that BLI provides a significant improvement in vehicle performance if the inlet distortion can be controlled, thus encouraging the pursuit of active flow control (AFC) as a BLI enabling technology. The effectiveness of active flow control in reducing engine inlet distortion was assessed using a 6% scale model of a 30% BLI offset, diffusing inlet. The experiment was conducted in the NASA Langley Basic Aerodynamics Research Tunnel with a model inlet designed specifically for this type of testing. High mass flow pulsing actuators provided the active flow control. Measurements were made of the onset boundary layer, the duct surface static pressures, and the mass flow through the duct and the actuators. The distortion was determined by 120 total pressure measurements located at the aerodynamic interface plane. The test matrix was limited to a maximum freestream Mach number of 0.15 with scaled mass flows through the inlet for that condition. The data show that the pulsed actuation can reduce distortion from 29% to 4.6% as measured by the circumferential distortion descriptor DC60 using less than 1% of inlet mass flow. Closed loop control of the actuation was also demonstrated using a sidewall surface static pressure as the response sensor.

Gorton, Susan Althoff↗
NASA NTRS1992

Internal Shock Interactions in Propulsion/Airframe Integrated Three-Dimensional Sidewall Compression Scramjet Inlets

The advantages and design requirements of propulsion/airframe integration for high Mach number flight have led to extensive study of the three-dimensional sidewall compression scramjet inlet in recent years. Recent research publications have indicated testing over a broad range of Mach number (2 to 18) in a variety of test gases, such as air, helium, and tetrafluoromethane. Multiple experimental techniques have been employed to obtain detailed internal shock interaction data, performance data, and inlet starting limits. Computational fluid dynamics has been effectively used for preliminary parametric studies as well as in parallel with experiments to aid in the explanation of unusual or unexpected flow phenomena. Inlets of this genre afford a relatively simple, generic geometry while producing a highly complex, three-dimensional flow field dominated by shock/shock and shock/boundary layer interactions. While the importance of the viscous effects in high speed inlet interactions is recognized, the present work addresses in a parametric fashion the inviscid effects of leading edge sweep, sidewall compression, and inflow Mach number on the internal shock structure in terms of inlet compression and mass capture. In the process, the source of the of the Mach number invariance with leading edge sweep for a constant sidewall compression class of inlet is identified, and a previously undocumented spillage phenomenon in a constant effective wedge angle class of inlets is discussed.

Holland, Scott D.↗
NASA NTRS2007

TBCC Inlet Experiments and Analysis

A research plan is being implemented at NASA to investigate inlet mode transition for turbine-based combined-cycle (TBCC) propulsion for the hypersonic community. Unresolved issues have remained on how to design an inlet system to supply both a turbine engine and a ram/scramjet flowpath that operate with both high performance and stability. The current plan is aimed at characterizing the design, performance and operability of TBCC inlets through a series of experiments and analyses. A TBCC inlet has been designed that is capable of high performance (near MIL-E-5008B recovery) with smooth transitioning characteristics. Traditional design techniques were used in an innovative approach to balance the aerodynamic and mechanical constraints to create a new TBCC inlet concept. The inlet was designed for top-end Mach 7 scramjet speeds with an over/under turbine that becomes cocooned beyond its Mach 4 peak design point. Conceptually, this propulsion system was picked to meet the needs of the first stage of a two-stage to orbit vehicle. A series of increasing fidelity CFD-based tools are being used throughout this effort. A small-scale inlet experiment is on-going in the GRC 1'x1' Supersonic Wind Tunnel (SWT). Initial results from both the CFD analyses and test are discussed showing that high performance and smooth mode transitions are possible. The effort validates the design and is contributing to a large-scale inlet/propulsion test being planned for the GRC 10' x10' SWT. This large-scale effort provide the basis for a Combined Cycle Engine Testbed, (CCET), that will be able to address integrated propulsion system and controls objectives.

Saunders, Dave↗
NASA NTRS2008

Forebody and Inlet Design for the HIFiRE 2 Flight Test

A forebody and inlet have been designed for the HIFiRE 2 scramjet flight test. The test will explore the operating, performance, and stability characteristics of a simple hydrocarbon-fueled scramjet combustor as it transitions from dual-mode to scramjet-mode operation and during supersonic combustion at Mach 8+ flight conditions. Requirements for the compression system were derived from inlet starting and combustor inflow requirements as well as physical size constraints. The design process is described. A planar, fixed geometry, mixed compression concept was used to produce laterally uniform flow at the inlet entrance and a conservative amount of internal contraction with respect to inlet starting. A grid sensitivity study was performed so that important flow physics caused by three-dimensional shock boundary layer interactions could be captured with confidence. Results from low Mach number operability studies, nominal trajectory cases, and high dynamic pressure heat load cases are discussed. The forebody and inlet solutions provide information for on-going combustor calculations, mass capture across the trajectory for fuel system design, and surface heating rates for thermal/structural analysis. The design has a one freestream Mach number margin for inlet starting, exceeds the high Mach number combustor entrance pressure requirement, produces high quality flow at the inlet exit for all Mach numbers and vehicle attitudes in the design space, and fits inside the booster shroud.

Ferlemann, Paul G.↗