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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 181 records · Page 10
All sourcesNASA NTRSDOE OSTI
NASA NTRS1978

Gas Path Sealing in Turbine Engines

A survey of gas path seals is presented with particular attention given to sealing clearance effects on engine component efficiency. The effects on compressor pressure ratio and stall margin are pointed out. Various case-rotor relative displacements, which affect gas path seal clearances, are identified. Forces produced by nonuniform sealing clearances and their effect on rotor stability are discussed qualitatively, and recent work on turbine-blade-tip sealing for high temperature is described. The need for active clearance control and for engine structural analysis is discussed. The functions of the internal-flow system and its seals are reviewed.

Ludwig, L. P.↗
NASA NTRS1978

Gas path sealing in turbine engines

Survey of gas path seals is presented with particular attention given to sealing clearance effects on engine component efficiency. The effects on compressor pressure ratio and stall margin are pointed out. Various case-rotor relative displacements, which affect gas path seal clearances, are identified. Forces produced by nonuniform sealing clearances and their effect on rotor stability are discussed qualitatively, and recent work on turbine-blade-tip sealing for high temperatures is described. The need for active clearance control and for engine structural analysis is discussed. The functions of the internal-flow system and its seals are reviewed.

Ludwig, L. P.↗
NASA NTRS1978

Gas path sealing in turbine engines

Gas path seals are discussed with emphasis on sealing clearance effects on engine component efficiency, compressor pressure ratio, and stall margin. Various case-rotor relative displacements, which affect gas path seal clearances, are identified. Forces produced by nonuniform sealing clearances and their effect on rotor stability are examined qualitatively, and recent work on turbine-blade-tip sealing for high temperatures is described. The need for active clearance control and for engine structural analysis is discussed. The functions of the internal-flow system and its seals are reviewed.

Ludwig, L. P.↗
NASA NTRS1980

Expanded study of feasibility of measuring in-flight 747/JT9D loads, performance, clearance, and thermal data

The JT9D jet engine exhibits a TSFC loss of about 1 percent in the initial 50 flight cycles of a new engine. These early losses are caused by seal-wear induced opening of running clearances in the engine gas path. The causes of this seal wear have been identified as flight induced loads which deflect the engine cases and rotors, causing the rotating blades to rub against the seal surfaces, producing permanent clearance changes. The real level of flight loads encountered during airplane acceptance testing and revenue service and the engine's response in the dynamic flight environment were investigated. The feasibility of direct measurement of these flight loads and their effects by concurrent measurement of 747/JT9D propulsion system aerodynamic and inertia loads and the critical engine clearance and performance changes during 747 flight and ground operations was evaluated. A number of technical options were examined in relation to the total estimated program cost to facilitate selection of the most cost effective option. It is concluded that a flight test program meeting the overall objective of determining the levels of aerodynamic and inertia load levels to which the engine is exposed during the initial flight acceptance test and normal flight maneuvers is feasible and desirable. A specific recommended flight test program, based on the evaluation of cost effectiveness, is defined.

Sallee, G. P.↗
NASA NTRS1986

Heat transfer in shrouded rectangular cavities

Heat transfer with turbulent flow over shrouded rectangular cavities are numerically investigated. The geometry studied models flow through the clearance gap at the grooved tip of an axial turbine blade, where the blade rotates in close proximity to a stationary outer ring or shroud. The direction of relative shroud motion is always in opposition to the direction of the gas flow across the blade tip. Heat transfer characteristics and flow pattern in a cavity are found to be strongly influenced by the dimension of gap clearance, cavity geometry, and relative shroud movement.

Chyu, M. K.↗
NASA NTRS2008

Unsteady Analysis of Blade and Tip Heat Transfer as Influenced by the Upstream Momentum and Thermal Wakes

The effect of the upstream wake on the blade heat transfer has been numerically examined. The geometry and the flow conditions of the first stage turbine blade of GE s E3 engine with a tip clearance equal to 2 percent of the span was utilized. Based on numerical calculations of the vane, a set of wake boundary conditions were approximated, which were subsequently imposed upon the downstream blade. This set consisted of the momentum and thermal wakes as well as the variation in modeled turbulence quantities of turbulence intensity and the length scale. Using a one-blade periodic domain, the distributions of unsteady heat transfer rate on the turbine blade and its tip, as affected by the wake, were determined. Such heat transfer coefficient distribution was computed using the wall heat flux and the adiabatic wall temperature to desensitize the heat transfer coefficient to the wall temperature. For the determination of the wall heat flux and the adiabatic wall temperatures, two sets of computations were required. The results were used in a phase-locked manner to compute the unsteady or steady heat transfer coefficients. It has been found that the unsteady wake has some effect on the distribution of the time averaged heat transfer coefficient on the blade and that this distribution is different from the distribution that is obtainable from a steady computation. This difference was found to be as large as 20 percent of the average heat transfer on the blade surface. On the tip surface, this difference is comparatively smaller and can be as large as four percent of the average.

Ameri, Ali A.↗
NASA NTRS2017

A Parametric Study of Actuator Requirements for Active Turbine Tip Clearance Control of a Modern High Bypass Turbofan Engine

The efficiency of aircraft gas turbine engines is sensitive to the distance between the tips of its turbine blades and its shroud, which serves as its containment structure. Maintaining tighter clearance between these components has been shown to increase turbine efficiency, increase fuel efficiency, and reduce the turbine inlet temperature, and this correlates to a longer time-on-wing for the engine. Therefore, there is a desire to maintain a tight clearance in the turbine, which requires fast response active clearance control. Fast response active tip clearance control will require an actuator to modify the physical or effective tip clearance in the turbine. This paper evaluates the requirements of a generic active turbine tip clearance actuator for a modern commercial aircraft engine using the Commercial Modular Aero-Propulsion System Simulation 40k (C-MAPSS40k) software that has previously been integrated with a dynamic tip clearance model. A parametric study was performed in an attempt to evaluate requirements for control actuators in terms of bandwidth, rate limits, saturation limits, and deadband. Constraints on the weight of the actuation system and some considerations as to the force which the actuator must be capable of exerting and maintaining are also investigated. From the results, the relevant range of the evaluated actuator parameters can be extracted. Some additional discussion is provided on the challenges posed by the tip clearance control problem and the implications for future small core aircraft engines.

Turbine↗
NASA NTRS1997

Laser Anemometer Measurements of the Flow Field in a 4:1 Pressure Ratio Centrifugal Impeller

A laser-doppler anemometer was used to obtain flow-field velocity measurements in a 4:1 pressure ratio, 4.54 kg/s (10 lbm/s), centrifugal impeller, with splitter blades and backsweep, which was configured with a vaneless diffuser. Measured through-flow velocities are reported for ten quasi-orthogonal survey planes at locations ranging from 1% to 99% of main blade chord. Measured through-flow velocities are compared to those predicted by a 3-D viscous steady flow analysis (Dawes) code. The measurements show the development and progression through the impeller and vaneless diffuser of a through-flow velocity deficit which results from the tip clearance flow and accumulation of low momentum fluid centrifuged from the blade and hub surfaces. Flow traces from the CFD analysis show the origin of this deficit which begins to grow in the inlet region of the impeller where it is first detected near the suction surface side of the passage. It then moves toward the pressure side of the channel, due to the movement of tip clearance flow across the impeller passage, where it is cut by the splitter blade leading edge. As blade loading increases toward the rear of the channel the deficit region is driven back toward the suction surface by the cross-passage pressure gradient. There is no evidence of a large wake region that might result from flow separation and the impeller efficiency is relatively high. The flow field in this impeller is quite similar to that documented previously by NASA Lewis in a large low-speed backswept impeller.

Skoch, G. J.↗
NASA NTRS1985

Dual clearance squeeze film damper

A dual clearance hydrodynamic liquid squeeze film damper for a gas turbine engine is described. Under normal operating conditions, the device functions as a conventional squeeze film damper, using only one of its oil films. When an unbalance reaches abusive levels, as may occur with a blade loss or foreign object damage, a second, larger clearance film becomes active, controlling vibration amplitudes in a near optimum manner until the engine can be safely shut down and repaired.

Fleming, D. P.↗
NASA NTRS1978

Effect of rotor tip clearance and configuration on overall performance of a 12.77-centimeter tip diameter axial-flow turbine

The rotor tip clearance was obtained by use of a recess in the casing above the rotor blades and also by use of a reduced blade height. For the recessed casing configuration, the optimum rotor blade height was found to be the one where the rotor tip diameter was equal to the stator tip diameter. The tip clearance loss associated with this optimum recessed casing configuration was less than that for the reduced blade height configuration.

Haas, J. E.↗
NASA NTRS1979

Effect of rotor tip clearance and configuration on overall performance of a 12.77-centimeter tip diameter axial-flow turbine

An extensive experimental investigation was made to determine the effect of varying the rotor tip clearance of a 12.77-cm-tip diameter, single-stage, axial-flow reaction turbine. In this investigation, the rotor tip clearance was obtained by use of a recess in the casing above the rotor blades and also by use of a reduced blade height. For the recessed casing configuration, the optimum rotor blade height was found to be the one where the rotor tip diameter was equal to the stator tip diameter. The tip clearance loss associated with this optimum recessed casing configuration was less than that for the reduced blade height configuration.

Haas, J. E.↗
NASA NTRS1999

Heat Transfer and Flow on the First Stage Blade Tip of a Power Generation Gas Turbine: Experimental Results - Part 1

A combined computational and experimental study has been performed to investigate the detailed distribution of convective heat transfer coefficients on the first stage blade tip surface for a geometry typical of large power generation turbines(>100MW). This paper is concerned with the design and execution of the experimental portion of the study. A stationary blade cascade experiment has been run consisting of three airfoils, the center airfoil having a variable tip gap clearance. The airfoil models the aerodynamic tip section of a high pressure turbine blade with inlet Mach number of 0.30, exit Mach number of 0.75, pressure ratio of 1.45, exit Reynolds number based on axial chord of 2.57 x 10(exp 6), and total turning of about 110 degrees. A hue detection based liquid crystal method is used to obtain the detailed heat transfer coefficient distribution on the blade tip surface for flat, smooth tip surfaces with both sharp and rounded edges. The cascade inlet turbulence intensity level took on values of either 5% or 9%. The cascade also models the casing recess in the shroud surface ahead of the blade. Experimental results are shown for the pressure distribution measurements on the airfoil near the tip gap, on the blade tip surface, and on the opposite shroud surface. Tip surface heat transfer coefficient distributions are shown for sharp-edge and rounded-edge tip geometries at each of the inlet turbulence intensity levels.

Bunker, Ronald S.↗
NASA NTRS2003

System for Centering a Turbofan in a Nacelle During Tests

A feedback position-control system has been developed for maintaining the concentricity of a turbofan with respect to a nacelle during acoustic and flow tests in a wind tunnel. The system is needed for the following reasons: Thermal and thrust loads can displace the fan relative to the nacelle; In the particular test apparatus (see Figure 1), denoted as a rotor-only nacelle (RAN), the struts, vanes, and other stator components of a turbofan engine that ordinarily maintain the required concentricity in the face of thermal and thrust loads are not present; and The struts and stator components are not present because it is necessary to provide a flow path that is acoustically clean in the sense that the measured noise can be attributed to the fan alone. The system is depicted schematically in Figure 2. The nacelle is supported by two struts attached to a two-axis traverse table located outside the wind-tunnel wall. Two servomotors acting through 100:1 gearboxes drive the table along the Y and Z axes, which are perpendicular to the axis of rotation. The Y and Z components of the deviation from concentricity are measured by four laser displacement sensors mounted on the nacelle and aimed at reflective targets on the center body, which is part of the fan assembly. The outputs of the laser displacement sensors are digitized and processed through a personal computer programmed with control software. The control output of the computer commands the servomotors to move the table as needed to restore concentricity. Numerous software and hardware travel limits and alarms are provided to maximize safety. A highly ablative rub strip in the nacelle minimizes the probability of damage in the event that a deviation from concentricity exceeds the radial clearance [<0.004 in. (<0.1 mm)] between the inner surface of the nacelle and the tips of the fan blades. To be able to prevent an excursion in excess of the tip clearance, the system must be accurate enough to control X and Y displacements to within 0.001 in. (.0.025 mm). One characteristic essential to such accuracy is sufficient rigidity in the mechanical components of the system to prevent excitation of vibrations in the strut/ nacelle subsystem. The need for such a high degree of accuracy prompted a comprehensive analysis of sources of measurement and control errors, followed by rigorous design efforts to minimize these errors. As a result, the design of the system incorporates numerous improvements in hardware, software, and operational procedures.

Cunningham, Cameron C.↗
NASA NTRS1992

TCGRID: A three dimensional C-grid generator for turbomachinery

A fast 3-D grid code for turbomachinery was developed. The code, TCGRID (Turbomachinery C-GRID), can generate either C or H type grids for fairly arbitrary axial or radical turbomachinery geometries. The code also has limited blocked grid capability and can generate an axisymmetric H type grid upstream of the blade row or an O type grid within the tip clearance region. Hub and tip geometries are input as a simple list of pairs. All geometric data is handled using parametric splines so that geometries that turn 90 degrees can be handled without difficulty. Blade input is in standard MERIDL or Lewis compressor design code format. TCGRID adds leading and trailing edge circles to MERIDL geometries and intersects the blade with the hub and tip if necessary using a novel intersection algorithm. The procedure used to generate the grid is given. Output is in PLOT3D format, which can also be read by the RVC3D (Rotor Viscous Code 3-D) Navier-Stokes code for turbomachinery. Intermediate 2-D or 3-D grids useful for debug and other purposes can also be output using a convenient output flag. A grid generated figure is given.

Chima, Rodrick V.↗
NASA NTRS1993

Supersonic flow and shock formation in turbine tip gaps

Shock formation due to overexpansion of supersonic flow at the inlet to the tip clearance gap of a turbomachine has been studied. As the flow enters the tip gap, it accelerates around the blade pressure-side corner creating a region of minimum static pressure. The 'free streamline' separates from the wall at the corner; and, for Mach numbers greater than about 1.3, it curves back to intersect the blade tip. At this point, the freestream flow is abruptly turned parallel to the surface, giving rise to an oblique shock. The results are consistent with compressible sharp-edged orifice flow calculations found in the literature and with the theory of oblique shock wave formation in supersonic flow over a wedge. For freestream Mach numbers of 1.4 to 1.8, wave angles are 43 to 54 deg, and turning angles are 9 to 20 deg; as the Mach number increases, the angle of turn also increases. It appears that in a turbine, after separating from the inlet corner, the flow reattaches on the blade tip and an oblique shock is formed at 0.4-1.4 tip gap heights into the clearance gap. The resulting shock-boundary layer interaction may contribute to further enhancement of already high heat transfer to the blade tip in this region. This in turn could lead to higher blade temperatures and adversely affect blade life and turbine efficiency.

Moore, John↗
NASA NTRS1985

Partial rotor-to-stator rub demonstration

A rotor radial rub typically occurs in seals or at a blade tip or shroud when there is insufficient clearance, high vibration, or the shaft equilibrium position has been displaced to effectively limit the clearance (eccentricity). There are two extreme cases of radial rubs: full annular rub, when the rotor maintains continuous contact with the seal, etc.; and a partial rub, when the contact occurs during a fraction of the precession period. They both involve similar physical phenomena such as friction and modification of stiffness. In partial rubs with consecutive impacts, a significant average value of radial force is generated. This results in shaft average displacement in the direction opposite the rub location. The rotor rig demonstrates the characteristics of a partial lateral rub of varying severity and location. These characteristics include: (1) subharmonic components as a function of rotative speed/first balance resonance ratio and radial force; (2) higher harmonic content as a function of severity; (3) increased average rotor stiffness resulting in increased first balance resonance speed; and (4) change in overall orbital pattern as a sum of the unbalance response (1x) and subharmonic response (1nx).

Grissom, R.↗
NASA NTRS1991

The role of tip clearance in high-speed fan stall

A numerical experiment has been carried out to define the near-stall casing endwall flowfield of a high-speed fan rotor. The experiment used a simulation code incorporating a simple clearance model, whose calibration is presented. The results of the simulation show that the interaction of the tip leakage vortex and the in-pasage shock plays a major role in determining the fan flow range. More specifically, the computations imply that it is the area increase of this vortex as it passes through the in-passage shock, which is the source of the blockage associated with stall. In addition, for fans of this type, it is the clearance over the forward portion of the fan blade which controls the flow processes leading to stall.

Adamczyk, J. J.↗
NASA NTRS1950

Free-Space Oscillating Pressures Near the Tips of Rotating Propellers

The theory is given for calculating the free-space oscillating pressures associated with a rotating propeller, at any point in space. Because of its complexity this analysis is convenient only for use in the critical region near the propeller tips where the assumptions used by Gutin to simplify his final equations are not valid. Good agreement was found between analytical and experimental results in the tip Mach number range 0.45 to two, three, four, five, six, on eight-blade propellers and for a range of tip clearances from 0.04 to 0.30 times the propeller diameter. If the power coefficient, tip Mach number, and the tip clearance are known for a given propeller, the designer may determine from these charts the average maximum free-space oscillating pressure in the critical region near the plane of rotation. A section of the report is devoted to the fuselage response to these oscillating pressures and indicates some of the factors to be considered in solving the problems of fuselage vibration and noise.

Hubbard, Harvey H↗