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At least 127 records · Page 7

All sources NASA NTRS DOE OSTI

NASA NTRS1977

Aeroacoustic performance of a scoop inlet

Results of a low speed wind tunnel test program are presented which demonstrate the aerodynamic and acoustic performance of a scoop inlet. Engine noise that would normally propagate toward the ground is directed upward by the extended lower lip of the scoop inlet. In addition, more of the scoop airflow comes in from above the inlet than below, leading to relatively higher surface velocities on the upper lip and lower surface velocities on the lower lip. These lower velocities on the lower lip result in a higher attainable angle of attack before internal flow separation occurs.

Abbott, J. M.↗

NASA NTRS1978

Effect of design changes on aerodynamic and acoustic performance of translating-centerbody sonic inlets

An experimental investigation was conducted to determine the effect of design changes on the aerodynamic and acoustic performance of translating centerbody sonic inlets. Scale model inlets were tested in the Lewis Research Center's V/STOL wind tunnel. The effects of centerbody position, entry lip contraction ratio, diffuser length, and diffuser area ratio on inlet total pressure recovery, distortion, and noise suppression were investigated at static conditions and at forward velocity and angle of attack. With the centerbody in the takeoff position (retracted), good aerodynamic and acoustic performance was attained at static conditions and at forward velocity. At 0 deg incidence angle with a sound pressure level reduction of 20 dB, the total pressure recovery was 0.986. Pressure recovery at 50 deg was 0.981. With the centerbody in the approach position (extended), diffuser flow separation occurred at an incidence angle of approximately 20 deg. However, good performance was attained at lower angles. With the centerbody in the takeoff position the ability of the inlet to tolerate high incidence angles was improved by increasing the lip contraction ratio. However, at static conditions with the centerbody in the approach position, an optimum lip contraction ratio appears to exist, with both thinner and thicker lips yielding reduced performance.

Miller, B. A.↗

NASA NTRS1991

The spatial characteristics of the B helical mode for a choked axisymmetric nozzle

An acoustic near-field study was performed for an axisymmetric conical nozzle operated at a fully expanded Mach number of 1.3. The acoustic measurements were performed in the nozzle exit plane using multiple sensors arranged around the periphery of the nozzle. The acquired data were simultaneously digitized. The acoustic spatial characteristics, describing the large scale structure associated with the preferred shear layer instability mode, were determined for the dominant B screech mode. The nozzle was fitted with a lip thickening device to determine the effect of this geometric variable on the spatial structure of the jet. For the thin-lipped configuration, the flapping structure of the B screech mode was found to precess in a time-dependent manner about the jet axis. An increasing of the nozzle exit lip thickness altered the spatial characteristics of this mode from a flapping to a time-dependent flapping or spinning. Using the auto-bicoherence spectrum, frequency dependencies were found to exist in the acoustic data of the thick-lipped configuration. These dependencies may be related to the nonlinearity of the jet and/or the time-dependent nature of the B mode spatial structure.

Ponton, Michael K.↗

NASA NTRS1992

Application of computational fluid dynamics to the design of the film cooled STME subscale nozzle for the National Launch System

The status of computational fluid dynamics (CFD) calculations for the Space Transportation Main Engine (STME) film/dump cooled nozzle design is presented, with an emphasis on the timely impact of CFD on the design of the sub-scale nozzle coolant system. The following aspects of the sub-scale coolant delivery system were analyzed with CFD: 1) a design trade study of a mechanical flow splitting device for uniform distribution of the subsonic cavity flow, 2) a design trade study of the subsonic cavity lip to achieve film integrity, and 3) an analysis of the primary flow interaction with the core/secondary coolant streams. All design calculations were performed with the Generalized Aerodynamic Simulation Program (GASP), a 3-D, multi-block, generalized Navier-Stokes code capable of solving with frozen, finite-rate or equilibrium chemical kinetics. The initial design of the subsonic cavity flow used square posts to distribute the sonic orifice jets into a uniform flow. Calculations for this design indicated that an unacceptable mal-distribution of film occurred. Design modifications involving curved and slotted posts were computed in an effort to uniformly distribute the secondary coolant flow. Analysis of these configurations showed that although the flowfield improved in uniformity, it was still unacceptable, especially at higher feed pressures. Results from these studies were then incorporated into a design that resulted in the insertion of a porous metal ring into the subsonic cavity. Subsequent water flow model studies showed that this concept was successful in uniformly distributing flow exiting the cavity. In addition to the design of the subsonic cavity, CFD was also used to analyze the secondary coolant lip and the primary flow interaction with the core/secondary coolant streams. A series of calculations were first performed to modify the subsonic cavity lip contour. The flow over the modified lip was then computed simultaneously with the primary injectors to determine the impact of the subsonic coolant stream on the primary slot jets. Pressure, temperature, velocity, and coolant mass fraction contours are presented for these configurations. Information is given in viewgraph form.

Garrett, Joseph L.↗

NASA NTRS1992

Shock interference prediction using direct simulation Monte Carlo

The shock interaction produced when an incident shock impinges on an inlet cowl lip is investigated for a cowl lip radius of 0.1 in. and flight conditions of Mach 15 and 35-km altitude. This problem is of interest in the design of hypersonic flight vehicles because the interference heating at the cowl lip surface from such a shock interaction is expected to be very high and the design of adequate thermal protection in this region will be challenging. Noncontinuum effects may be significant at this combination of altitude and small dimensions. If this is so, a computational solution based on the Navier-Stokes equations could significantly overpredict the surface heating on the cowl lip. Therefore, a kinetic theory solution technique, direct simulation Monte Carlo (DSMC), is used to predict the flowfield and surface heating. The details of modeling this problem and some preliminary results are presented. The same qualitative flow patterns are seen with DSMC and a continuum solution of the shock interference case, but the peak heating predicted with DSMC is somewhat lower.

Carlson, Ann B.↗

NASA NTRS1995

Acoustic Study of B Helical Mode for Choked Axisymmetric Nozzle

An acoustic near-field study was performed for an axisymmetric conical convergent nozzle operating at a pressure ratio corresponding to a fully expanded Mach number of 1.3. The acoustic measurements were performed in the nozzle exit plane using multiple sensors arranged around the periphery of the nozzle. The acquired data were simultaneously digitized. The acoustic spatial characteristics, describing the large-scale structure associated with the preferred shear layer instability mode, were determined for the dominant B screech mode. The nozzle was fitted with a lip thickening device to determine the effect of this geometric variable on the spatial structure of the jet. For the thin-lipped configuration, the flapping structure of the B screech mode was found to precess in a time- dependent manner about the jet axis. Increasing of the nozzle exit lip thickness altered the spatial characteristics of this mode from a flapping to a time-dependent flapping or spinning. Using the autobicoherence spectrum, frequency dependencies were found to exist in the acoustic data of the thick-lipped configuration. These dependencies may be related to the nonlinearity of the jet and/or the time-dependent nature of the B mode spatial structure.

Ponton, Michael K.↗

NASA NTRS1947

Tests of Submerged Duct Installation on the Ryan FR-1 Airplane in the Ames 40- by 80-Foot Wind Tunnel

An investigation of an NACA submerged intake installation on the Ryan FR-1 was conducted to determine the full-scale aerodynamic characteristics of this installation. In addition, tests were conducted on the submerged inlet with revised entrance lips and deflectors to determine the configuration which would result in the best dynamic pressure recovery measured at the inlet for this installation without a major rework of the entrance. Stalling of the air flow over the inner lip surface created excessive dynamic pressure losses with the original entrance. The revised entrance produced a 12-percent increase in dynamic pressure recovery at the design high-speed inlet-velocity ratio and resulted in an improvement of thte critical-speed characteristics of the entrance lip. A complete redesign of the entrance including a decrease in ramp angle and adjustment of lip camber is necessary to secure optimum results from this submerged duct installation.

Martin, Norman J.↗

NASA NTRS2008

Transient Three-Dimensional Startup Side Load Analysis of a Regeneratively Cooled Nozzle

The objective of this effort is to develop a computational methodology to capture the startup side load physics and to anchor the computed aerodynamic side loads with the available data from a regeneratively cooled, high-aspect-ratio nozzle, hot-fired at sea level. The computational methodology is based on an unstructured-grid, pressure-based, reacting flow computational fluid dynamics and heat transfer formulation, a transient 5 s inlet history based on an engine system simulation, and a wall temperature distribution to reflect the effect of regenerative cooling. To understand the effect of regenerative wall cooling, two transient computations were performed using the boundary conditions of adiabatic and cooled walls, respectively. The results show that three types of shock evolution are responsible for side loads: generation of combustion wave; transitions among free-shock separation, restricted-shock separation, and simultaneous free-shock and restricted shock separations; along with the pulsation of shocks across the lip, although the combustion wave is commonly eliminated with the sparklers during actual test. The test measured two side load events: a secondary and lower side load, followed by a primary and peak side load. Results from both wall boundary conditions captured the free-shock separation to restricted-shock separation transition with computed side loads matching the measured secondary side load. For the primary side load, the cooled wall transient produced restricted-shock pulsation across the nozzle lip with peak side load matching that of the test, while the adiabatic wall transient captured shock transitions and free-shock pulsation across the lip with computed peak side load 50% lower than that of the measurement. The computed dominant pulsation frequency of the cooled wall nozzle agrees with that of a separate test, while that of the adiabatic wall nozzle is more than 50% lower than that of the measurement. The computed teepee-like formation and the tangential motion of the shocks during lip pulsation also qualitatively agree with those of test observations. Moreover, a third transient computation was performed with a proportionately shortened 1 s sequence, and lower side loads were obtained with the higher ramp rate.

Wang, Ten-See↗

NASA NTRS2011

A Zero-Gravity Cup for Drinking Beverages in Microgravity

To date, the method for astronauts to drink liquids in microgravity or weightless environments is to suck the liquid from a bag or pouch through a straw. A new beverage cup works in microgravity and allows astronauts to drink liquids from a cup in a manner consistent with that on Earth. The cup is capable of holding beverages with an angled channel running along the wall from the bottom to the lip. In microgravity, a beverage is placed into the cup using the galley dispenser. The angled channel acts as an open passage that contains only two sides where capillary forces move the liquid along the channel until it reaches the top lip where the forces reach an equilibrium and the flow stops. When one sips the liquid at the lip of the channel, the capillary force equilibrium is upset and more liquid flows to the lip from the reservoir at the bottom to re-establish the equilibrium. This sipping process can continue until the total liquid contents of the cup is consumed, leaving only a few residual drops about the same quantity as in a ceramic cup when it is drunk dry on Earth.

Pettit, Donald R.↗

NASA NTRS2019

Scale–Resolving Simulations of a Fundamental Trailing–Edge Cooling Slot Using a Discontinuous–Galerkin Spectral–Element Method

The accurate prediction of turbulent mixing in high-pressure turbines that incorporate various airfoil surface-cooling strategies is becoming increasing critical to the design of modern gas turbine engines where the quest for improved efficiency is driving compressor overall pressure ratios and turbine inlet temperatures to much higher levels than ever before. In the present paper, a recently developed computational capability for accurate and efficient scale–resolving simulations of turbomachinery is extended to study the turbulent mixing mechanism of a simplified abstraction of an airfoil trailing-edge cooling slot - a plane wall jet with finite lip thickness discharging into an ambient flow. The computational capability is based on an entropy– stable, discontinuous–Galerkin approach that extends to arbitrarily high orders of spatial and temporal accuracy. The numerical results show that the present simulations capture the trends observed in the experiments. Discrepancies between the simulations and experiments are believed to be due to differences in the inflow profiles and tunnel side–wall effects. The thick lip configuration leads to a thicker wake and higher unsteadiness in the wall jet compared to the thin lip. A detailed comparison of the turbulent flowfields is presented to highlight differences arising due to lip thickness variations.

Scale Resolving↗

NASA NTRS1975

Aeroacoustic performance of scale model sonic inlets

A low speed wind tunnel was used to evaluate the inflight aeroacoustic performance of several single- and multiple-passage sonic inlets. Takeoff and approach geometries were tested, and the effects of inlet lip and diffuser design were determined. Results indicate that the single passage geometries, in particular a cylindrical centerbody takeoff geometry and a bulb shaped centerbody approach geometry, provide the highest level of aeroacoustic performance. Increasing inlet lip contraction ratio extends the maximum incidence angle for attached lip flow, while increasing inlet diffuser length results in higher total pressure recovery for a given amount of noise suppression.

Abbott, J. M.↗

NASA NTRS1975

Vortex shedding from a blunt trailing edge with equal and unequal external mean velocities

A flow visualization study showed that strong Karman vortices are developed behind the blunt trailing edge of a plate when the free stream velocities over both surfaces are equal. These vortices tend to disappear when the surface velocities are unequal. This observation provided an explanation for the occurrence and disappearance of the lip noise often present in coaxial jets. Vortex formation and lip noise occurred at a Strouhal number of about 0.2 based on the lip thickness and the average of the external steady-state velocities. Results from theoretical calculations of the vortex formation, based on an inviscid, incompressible analysis of the motion of point vortices, were in good agreement with the experimental observations.

Brinich, P. F.↗

NASA NTRS1975

Aeroacoustic performance of scale model sonic inlets

Various means can be used to vary engine inlet throat areas in order to reduce noise generated at takeoff and approach. Low-speed wind-tunnel tests were performed to evaluate the aeroacoustic performance of two takeoff geometries (cylindrical centerbody and bulb-shaped centerbody) and four approach geometries (bulb-shaped centerbody, annular ring, radial vanes, and step diffuser). The effects of inlet internal lip shape and diffuser length on sonic inlet performance were also investigated. It was found that the single-passage geometries, in particular a cylindrical centerbody takeoff geometry and a bulb-shaped centerbody approach geometry, provide the highest level of aeroacoustic performance. Increasing the inlet lip contraction ratio extends the maximum incidence angle for attached lip flow, while increasing the inlet diffuser length increases total pressure recovery for a given amount of noise suppression.

Abbott, J. M.↗

NASA NTRS1975

An investigation of several NACA 1-series inlets at Mach numbers from 0.4 to 1.29 for mass flow ratios near 1.0

An investigation to determine the performance of eight NACA 1-series inlets at massflow ratios near 1.0 was conducted in the Langley 16-foot transonic tunnel. The inlet diameter ratios (ratio of inlet diameter to maximum diameter) were 0.85 and 0.89 for an inlet length ratio (ratio of inlet length to maximum diameter) of 1.0. Inlet lip radius varied from 0.061 cm to 0.251 cm, and internal contraction area ratio (ratio of inlet area to throat area) varied from 1.006 to 1.201. Reynolds number based on model maximum diameter ranged from 3,600,000 at a Mach number of 400,000 to 5,900,000 at a Mach number of 1.29. The results indicate that nearly uniform pressure distributions on a given inlet were obtained over a limited range of mass-flow ratios and Mach numbers. When inlet lip thickness was increased by means of lip radius or contraction ratio, the inlet critical Mach number decreased. Drag-divergence Mach number inferred from forebody pressure integrations was above 0.94 for most of the inlets tested.

Re, R. J.↗

NASA NTRS1976

Low speed wind tunnel investigation of the aerodynamic and acoustic performance of several sonic inlet takeoff and approach geometries

A series of tests was conducted to determine the aerodynamic and acoustic performance of several sonic inlet takeoff and approach geometries. The effects of inlet lip shape and diffuser length were also investigated. The tests were conducted in a low-speed wind tunnel at free-stream velocities of 0 and 45 meters per second. Inlet incidence angle was varied from 0 deg to 50 deg. The inlets were sized to fit a 13.97-centimeter-diameter fan. In terms of the highest level of inlet total pressure recovery for a given amount of noise suppression, a cylindrical centerbody takeoff geometry and a bulb-shaped centerbody approach geometry provided the best results over all conditions of free-stream velocity and incidence angle. Increasing inlet lip contraction ratio extended the maximum incidence angle for attached lip flow, while increasing inlet diffuser length resulted in a higher total pressure recovery for a given amount of noise suppression.

Abbott, J. M.↗

NASA NTRS1978

Low speed test of a high-bypass-ratio propulsion system with an asymmetric inlet designed for a tilt-nacelle V/STOL airplane

A large scale model of a lift/cruise fan inlet designed for a tilt nacelle V/STOL airplane was tested with a high bypass ratio turbofan. Testing was conducted at low freestream velocities with inlet angles of attack ranging from 0 deg to 120 deg. The operating limits for the nacelle were found to be related to inlet boundary layer separation. Small separations originating in the inlet diffuser cause little or no performance degradation. However, at sufficiently severe freestream conditions the separation changes abruptly to a lip separation. This change is associated with a significant reduction in nacelle net thrust as well as a sharp increase in fan blade vibratory stresses. Consequently, the onset of lip separation is regarded as the nacelle operating limit. The test verified that the asymmetric inlet design will provide high performance and stable operation at the design forward speed and angle of attack conditions. At some of these, however, operation near the lower end of the design inlet airflow range is not feasible due to the occurrence of lip separation.

Syberg, J.↗

NASA NTRS1979

Self stabilizing sonic inlet

An inlet suitable for a turbine engine in a STOL, VTOL or CTOL aircraft is described. A circumferentially extended slot is provided in the inner surface of the air inlet at the windward side and downstream of the throat region. The slot communicates with a circumferential plenum chamber formed in the front of the air inlet just behind the lip. Circumferentially extending rows of apertures are provided on the lip establishing two sets of apertures spaced circumferentially away from the slot in opposite directions. The slot removes the boundary layer from the critical portion of the diffuser to minimize or eliminate flow pressure loss or separations resulting from diffusion or tuning. The apertures are in a region of low static pressure on the lip of the inlet and serve as a source of suction to cause air flow into the slot.

Miller, B. A.↗

NASA NTRS1981

Zero-length inlets for subsonic V/STOL aircraft

Zero-length inlet performance and associated fan blade stresses were determined during model tests in the NASA-LeRC 9-by 15-foot low-speed wind tunnel. The inlet models, which were installed on a 20-inch diameter fan unit, had different inlet lip contraction ratios as well as unslotted, slotted, and double slotted inlet lips. The inlet angle-of-attack boundaries for onset of flow separation were identified and compared to the operating requirements of several generically different subsonic V/STOL aircraft. The zero-length inlets, especially those with slotted lips, were able to satisfy these requirements without compromising the maximum cowl forebody radius. As an aid to the inlet design process, a unique relationship was established between the maximum surface Mach number associated with the separation boundary and the maximum-to-throat surface velocity ratio.

Glasgow, E. R.↗