Results for “AIRFLOW - WINGS, SLOTTED”

NASA NTRS1992

Positive Stop For Circulation-Control Slot

Rounded shoulder on bushing ensures repeatable stop position. Improved stop mechanism allows airflow slot on circulation-control airfoil to be limited to precise width. Shoulder bushing installed in duct so shoulder rests on inside surface of duct. Bolt slides in bushing in duct wall between more open and less open positions. No relative motion or wear between bolt and slot panel. Prevents bushing from being dislodged by stop forces and provides stiffer stopping surface than sleeve bushing bonded in duct wall.

Hunter, David↗

NASA NTRS1974

Forward slanted slot throat stability bypass to increase the stable airflow range of a Mach 2.5 inlet with 60 percent internal contraction

The results of an experimental investigation to increase the stable airflow operating range of a supersonic mixed-compression inlet are presented. Two forward-slanted slot stability-bypass entrance configurations were tested. In terms of diffuser-exit corrected airflow, a large inlet stable airflow range of 18.5 percent was obtained with the superior configuration if a constant pressure was maintained in the bypass plenum. Limited unstart angle-of-attack data are presented.

Shaw, R. J.↗

NASA NTRS1974

Throat stability-by pass systems to increase the stable airflow range of a Mach 2.5 inlet with 60-percent internal contraction

The results of an experimental investigation to increase the stable airflow range (without unstart) of a supersonic mixed-compression inlet are presented. Various stability bypass entrances were located on the cowl side of the inlet throat. The types of entrance were distributed porous (normal holes), forward-slanted slot, and distributed educated slots. A large stable airflow range was obtained for each entrance type if a constant pressure was maintained in the stability bypass plenum. The distributed porous entrance provided the largest stable airflow range. Inlet unstart angle of attack was unaffected by the entrances.

Mitchell, G. A.↗

NASA NTRS1985

Dynamic response characteristics of a circulation control rotor model pneumatic system

Numerical and experimental simulation of unsteady airflow through the control valve and slotted air duct of a circulation control rotor is described. The numerical analysis involves the solution of the quasi-one-dimensional compressible fluid-dynamic equations in the blade air duct together with the coupled isentropic flow equations for flow into the blade through the valve and out of the blade through the Coanda slot. Numerical solutions are compared with basic experimental results obtained for a mockup of a circulation control rotor and its pneumatic valving system. The pneumodynamic phenomena that were observed are discussed with particular emphasis on the characteristic system time lags associated with the response of the flow variables to transient and periodic control valve inputs.

Watkins, C. B.↗

NASA NTRS2004

Inlet Housing for a Partial-Admission Turbine

An inlet housing for a partial-admission turbine has been designed to cause the inlet airflow to make a smooth transition from an open circular inlet to an inlet slot. The smooth flow is required for purposes of measuring inlet flow characteristics and maximizing the efficiency of the turbine. A partial-admission turbine is a turbine in which the inlet slot occupies less than a complete circle around the rotor axis. In this case, the inlet slot occupies a 90 arc. The present special inlet-housing design is needed because the "bull nose" shape of a conventional turbine inlet housing fails to provide the required smooth transition in a partial-admission configuration and thereby gives rise to a loss of turbine efficiency and inaccuracies in inlet flow measurements. Upon entering the inlet housing through the circular opening, the flow encounters a "tongue"-shaped passageway, which serves as a ramp that diverts the flow to the first of two straight passages. This first passageway occupies a 90 arc and has a length equal to two passage heights. Instrumentation rakes for measuring the characteristics of the inlet flow are installed in this passageway. Just past the first straight passageway is the second one, which is narrower and leads to the 90 turbine inlet slot. This passageway is used to smooth the flow immediately prior to its passage through the turbine inlet slot. The length of this second passageway equals the length of the chord of a turbine vane. The inlet housing incorporates small ports for measuring static pressures at various locations of the flow, and incorporates bosses for the installation of the instrumentation rakes. The inlet housing also includes a flange at its inlet end for attachment to a circular inlet duct and a flange at its outlet end for attachment to the outer casing of the turbine.

Moye, Ralph↗

NASA NTRS1975

Effect of free-stream turbulence on film cooling

Film-cooling experiments were conducted at four levels of free-stream turbulence to test the hypothesis that the film-cooling effectiveness is inversely related to the free-stream turbulence level. The hot-gas operating conditions were held constant at a temperature of 590 K, a pressure of 1 atmosphere, and a velocity of 62 m/sec. The film-cooling air was at ambient inlet temperature, and the film-cooling flow rates were 2.5, 5.0, 7.5, and 10.0 percent of the total airflow. Blockage plates with blockage areas of 0, 52, 72, and 90 percent were placed upstream of the film-cooling slot and produced axial turbulence intensities of 7, 14, 23, and 35 percent, respectively. The film-cooling effectiveness decreased as much as 50 percent as the freestream turbulence intensity was increased from 7 to 35 percent. The value of the turbulent mixing coefficient used in previous work was compared with the axial turbulence intensity. The turbulent mixing coefficient was found to be 10 to 40 percent of the axial turbulence intensity.

Marek, C. J.↗

NASA NTRS1973

Throat-bypass bleed systems for increasing the stable airflow range of a Mach 2.50 axisymmetric inlet with 40-percent internal contraction

The results of an experimental investigation to increase the stable airflow range of a super sonic mixed-compression inlet are presented. Various throat-bypass bleeds were located on the inlet cowl. The bleed types were distributed porous normal holes, a forward slanted slot, or distributed educated slots. Large inlet stability margins were obtained with the inlet throat bleed systems if a constant pressure was maintained in the throat-bypass bleed plenum. Stability limits were determined for steady-state and limited transient internal air flow changes. Limited unstart angle-of-attack data are presented.

Sanders, B. W.↗

NASA NTRS1973

Improved design of a high response slotted plate overboard bypass valve for supersonic inlets

The electrohydraulically actuated slotted-plate bypass valve used to control the position of the normal shock during wind-tunnel investigations of supersonic inlets also has proven to be a valuable device for determining inlet dynamics and creating airflow disturbances. Operation of previous valves at high frequencies (to 100 Hz) for extended testing has resulted in numerous failures. An improved bypass-valve design is presented which increases the cyclic tolerance of the device considerably over past designs. The use of dynamic limit criteria to obtain an optimum actuator-piston size results in a frequency response which is flat within + or - 3 decibels to 120 Hz for a peak-to-peak variation of 20 percent of full area.

Webb, J. A., Jr.↗

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 NTRS2015

Investigation of Gas Seeding for Planar Laser-Induced Fluorescence in Hypersonic Boundary Layers

Numerical simulations of the gas-seeding strategies required for planar laser-induced fluorescence in a Mach 10 (approximately Mach 8.2 postshock) airflow were performed. The work was performed to understand and quantify the adverse effects associated with gas seeding and to assess various types of seed gas that could potentially be used in future experiments. In prior experiments, NO and NO2 were injected through a slot near the leading edge of a flatplate wedge model used in NASA Langley Research Center's 31 in. Mach 10 air tunnel facility. In this paper, nitric oxide, krypton, and iodine gases were simulated at various injection rates. Simulations showing the deflection of the velocity boundary layer for each of the cases are presented. Streamwise distributions of velocity and concentration boundary-layer thicknesses, as well as vertical distributions of velocity, temperature, and mass distributions, are presented for each of the cases. A comparison between simulated streamwise velocity profiles and experimentally obtained molecular tagging velocimetry profiles using a nitric oxide seeding strategy is performed to verify the influence of such a strategy on the boundary layer. The relative merits of the different seeding strategies are discussed. The results from a custom solver based on OpenFOAM version 2.2.1 are compared against results obtained from ANSYS® Fluent version 6.3.

Arisman, C. J.↗

NASA NTRS2001

Reaction Kernel Structure of a Slot Jet Diffusion Flame in Microgravity

Diffusion flame stabilization in normal earth gravity (1 g) has long been a fundamental research subject in combustion. Local flame-flow phenomena, including heat and species transport and chemical reactions, around the flame base in the vicinity of condensed surfaces control flame stabilization and fire spreading processes. Therefore, gravity plays an important role in the subject topic because buoyancy induces flow in the flame zone, thus increasing the convective (and diffusive) oxygen transport into the flame zone and, in turn, reaction rates. Recent computations show that a peak reactivity (heat-release or oxygen-consumption rate) spot, or reaction kernel, is formed in the flame base by back-diffusion and reactions of radical species in the incoming oxygen-abundant flow at relatively low temperatures (about 1550 K). Quasi-linear correlations were found between the peak heat-release or oxygen-consumption rate and the velocity at the reaction kernel for cases including both jet and flat-plate diffusion flames in airflow. The reaction kernel provides a stationary ignition source to incoming reactants, sustains combustion, and thus stabilizes the trailing diffusion flame. In a quiescent microgravity environment, no buoyancy-induced flow exits and thus purely diffusive transport controls the reaction rates. Flame stabilization mechanisms in such purely diffusion-controlled regime remain largely unstudied. Therefore, it will be a rigorous test for the reaction kernel correlation if it can be extended toward zero velocity conditions in the purely diffusion-controlled regime. The objectives of this study are to reveal the structure of the flame-stabilizing region of a two-dimensional (2D) laminar jet diffusion flame in microgravity and develop a unified diffusion flame stabilization mechanism. This paper reports the recent progress in the computation and experiment performed in microgravity.

Takahashi, F.↗

NASA NTRSDate not supplied

LES-Based Analysis of Film Cooling in a High-Pressure Turbine Rotor Including Effects of Purge and Tip Flow

Accurate modeling of coolant airflows, which form protective films over turbine blades, is essential for designing fuel-efficient and environmentally sustainable gas turbine engines. Excessive coolant reduces thermal efficiency, while insufficient coolant leads to blade overheating and causes damage. Therefore, precise prediction of flow field interactions with cooling air is critical for optimizing turbine performance. This study numerically investigates the cooling effectiveness of purge and film cooling flows within a high-pressure turbine (HPT) rotor using Large Eddy Simulation (LES). The study utilizes NASA Glenn Research Center’s Glenn-HT solver. The simulation models ethe conditions of the Penn State University START rotating rig. A high-fidelity structured mesh comprising up to 800 million cells is employed to resolve high-Reynolds number flow (Re ≈ 350,000) and to capture intricate secondary flow structures, including tip leakage and purge-induced vortices. Film cooling effectiveness computations are highly sensitive to boundary conditions at the cooling holes and to grid resolution. Even with well-resolved grids and included plena, strong mixing challenges traditional eddy viscosity models. To address this, a simplified configuration is simulated: a truncated row of shaped holes on the suction side near the leading edge and a row on the pressure side, both fed from internal plena while the purge slot and tip clearance are also modeled. Two isothermal LES cases are conducted at two distinct wall temperatures, which yield the adiabatic wall temperature and the heat transfer coefficient. The definition and means of computation of the effectiveness is discussed in this paper. The simulations reveal detailed three-dimensional unsteady flow features, including coherent vortical structures and secondary flows originating from the purge cavity. Film cooling effectiveness and Nusselt number distributions are presented for both the blade surface and tip, highlighting regions of elevated heat transfer and complex thermal behavior. These findings underscore the importance of high-resolution LES and realistic boundary conditions in capturing the dynamics of purge and film cooling, offering valuable insights for improving turbine blade design and thermal management strategies.

Secondary Flows↗

NASA NTRSDate not supplied

A GPU-based Approach for Turbomachinery Application

Accurate modeling of coolant airflows, which form protective films over turbine blades, is essential for designing fuel-efficient and environmentally sustainable gas turbine engines. Excessive coolant reduces thermal efficiency, while insufficient coolant leads to blade overheating and causes damage. Therefore, precise prediction of flow field interactions with cooling air is critical for optimizing turbine performance. This study numerically investigates the cooling effectiveness of purge and film cooling flows within a high-pressure turbine (HPT) rotor using Large Eddy Simulation (LES). The study utilizes NASA Glenn Research Center’s Glenn-HT solver. The simulation models ethe conditions of the Penn State University START rotating rig. A high-fidelity structured mesh comprising up to 800 million cells is employed to resolve high-Reynolds number flow (Re ≈ 350,000) and to capture intricate secondary flow structures, including tip leakage and purge-induced vortices. Film cooling effectiveness computations are highly sensitive to boundary conditions at the cooling holes and to grid resolution. Even with well-resolved grids and included plena, strong mixing challenges traditional eddy viscosity models. To address this, a simplified configuration is simulated: a truncated row of shaped holes on the suction side near the leading edge and a row on the pressure side, both fed from internal plena while the purge slot and tip clearance are also modeled. Two isothermal LES cases are conducted at two distinct wall temperatures, which yield the adiabatic wall temperature and the heat transfer coefficient. The definition and means of computation of the effectiveness is discussed in this paper. The simulations reveal detailed three-dimensional unsteady flow features, including coherent vortical structures and secondary flows originating from the purge cavity. Film cooling effectiveness and Nusselt number distributions are presented for both the blade surface and tip, highlighting regions of elevated heat transfer and complex thermal behavior. These findings underscore the importance of high-resolution LES and realistic boundary conditions in capturing the dynamics of purge and film cooling, offering valuable insights for improving turbine blade design and thermal management strategies.

Gas Turbine↗