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At least 1,009 records · Page 56
All sourcesNASA NTRSDOE OSTI
NASA NTRS1997

Performance and Vibratory Loads Data From a Wind-Tunnel Test of a Model Helicopter Main-Rotor Blade With a Paddle-Type Tip

An investigation was conducted in the Langley Transonic Dynamics Tunnel to obtain data to permit evaluation of paddle-type tip technology for possible use in future U.S. advanced rotor designs. Data was obtained for both a baseline main-rotor blade and a main-rotor blade with a paddle-type tip. The baseline and paddle-type tip blades were compared with regard to rotor performance, oscillatory pitch-link loads, and 4-per-rev vertical fixed-system loads. Data was obtained in hover and forward flight over a nominal range of advance ratios from 0.15 to 0.425. Results indicate that the paddle-type tip offers no performance improvements in either hover or forward flight. Pitch-link oscillatory loads for the paddle-type tip are higher than for the baseline blade, whereas 4-per-rev vertical fixed-system loads are generally lower.

Yeager, William T., Jr.↗
NASA NTRS1996

Unsteady Blade Row Interaction in a Transonic Turbine

Experimental data from jet-engine tests have indicated that unsteady blade row interaction effects can have a significant impact on the performance of multiple-stage turbines. The magnitude of blade row interaction is a function of both blade-count ratio and axial spacing. In the current research program, numerical simulations have been used to quantify the effects of blade count ratio on the performance of an advanced turbine geometries.

Dorney, Daniel J.↗
NASA NTRS1998

Heat Transfer on a Film-Cooled Blade - Effect of Hole Physics

A multi-block, three-dimensional Navier-Stokes code has been used to study the within-hole and near-hole physics in relation to heat transfer on a film-cooled blade. The flow domain consists of the coolant flow through the plenum and hole-pipes for the three staggered rows of shower-head holes on the VK1 rotor, and the main flow over the blade. A multi-block grid is generated that is nearly orthogonal to the various surfaces. It may be noted that for the VK1 rotor the shower-head holes are inclined at 30 deg. to the spanwise direction, and are normal to the streamwise direction on the blade. Wilcox's k-omega turbulence model is used. The present study provides a much better comparison for the heat transfer coefficient at the blade mid-span with the experimental data than an earlier analysis wherein coolant velocity and temperature distributions were specified at the hole exits rather than extending the computational domain into the hole-pipe and plenum. Details of the distributions of coolant velocity, temperature, k and omega at the hole exits are also presented.

Garg, Vijay K.↗
NASA NTRS1997

Effect of Velocity and Temperature Distribution at the Hole Exit on Film Cooling of Turbine Blades

An existing three-dimensional Navier-Stokes code (Arnone et al, 1991), modified Turbine Branch, to include film cooling considerations (Garg and Gaugler, 1994), has been used to study the effect of coolant velocity and temperature distribution at the hole exit on the heat transfer coefficient on three film-cooled turbine blades, namely, the C3X vane, the VKI rotor, and the ACE rotor. Results are also compared with the experimental data for all the blades. Moreover, Mayle's transition criterion (1991), Forest's model for augmentation of leading edge heat transfer due to free-stream turbulence (1977), and Crawford's model for augmentation of eddy viscosity due to film cooling (Crawford et al, 1980) are used. Use of Mayle's and Forest's models is relevant only for the ACE rotor due to the absence of showerhead cooling on this rotor. It is found that, in some cases, the effect of distribution of coolant velocity and temperature at the hole exit can be as much as 60 percent on the heat transfer coefficient at the blade suction surface, and 50 percent at the pressure surface. Also, different effects are observed on the pressure and suction surface depending upon the blade as well as upon the hole shape, conical or cylindrical.

Garg, V. K.↗
NASA NTRS1998

Rotor Blade Pressure Measurement in a Rotating Machinery Using Pressure and Temperature Sensitive Paints

Pressure and temperature sensitive paints have been utilized for the measurement of blade surface pressure and temperature distributions in a high speed axial compressor and an Allied Signal F109 gas turbine engine. Alternate blades were painted with temperature sensitive paints and then pressure sensitive paint. This combination allows temperature distributions to be accounted for when determining the blade suction surface pressure distribution. Measurements were taken and pressure maps on the suction surface of a blade were obtained over a range of rotational speeds. Pressure maps of the suction surface show-strong shock waves at the higher speeds.

Torgerson, S.↗
NASA NTRS1959

Experimental Evaluation of Cermet Turbine Stator Blades for Use at Elevated Gas Temperatures

The suitability of cermets for turbine stator blades of a modified turbojet engine was determined at an average turbine-inlet-gas temperature of 2000 F. Such an increase in temperature would yield a premium in thrust from a service engine. Because the cermet blades require no cooling, all the available compressor bleed air could be used to cool a turbine made from conventional ductile alloys. Cermet blades were first run in 100-hour endurance tests at normal gas temperatures in order to evaluate two methods for mounting them. The elevated gas-temperature test was then run using the method of support considered best for high-temperature operation. After 52 hours at 2000 F, one of the group of four cermet blades fractured probably because of end loads resulting from thermal distortion of the spacer band of the nozzle diaphragm. Improved design of a service engine would preclude this cause of premature failure.

Chiarito, Patrick T.↗
NASA NTRS1998

A Numerical Simulator for Three-Dimensional Flows Through Vibrating Blade Rows

The three-dimensional, multi-stage, unsteady, turbomachinery analysis, TURBO, has been extended to predict the aeroelastic and aeroacoustic response behaviors of a single blade row operating within a cylindrical annular duct. In particular, a blade vibration capability has been incorporated so that the TURBO analysis can be applied over a solution domain that deforms with a vibratory blade motion. Also, unsteady far-field conditions have been implemented to render the computational boundaries at inlet and exit transparent to outgoing unsteady disturbances. The modified TURBO analysis is applied herein to predict unsteady subsonic and transonic flows. The intent is to partially validate this nonlinear analysis for blade flutter applications, via numerical results for benchmark unsteady flows, and to demonstrate the analysis for a realistic fan rotor. For these purposes, we have considered unsteady subsonic flows through a 3D version of the 10th Standard Cascade, and unsteady transonic flows through the first stage rotor of the NASA Lewis, Rotor 67, two-stage fan.

Chuang, H. Andrew↗
NASA NTRS1999

Heat Transfer on a Film-Cooled Blade: Effect of Hole Physics

A multiblock, three-dimensional Navier Stokes code has been used to study the within-hole and near-hole physics in relation to heat transfer on a film-cooled blade. The flow domain consists of the coolant flow through the plenum and hole-pipes for the three staggered rows of shower-head holes on the VKI rotor, and the main flow over the blade. A multiblock grid is generated that is nearly orthogonal to the various surfaces. It may be noted that for the VKI rotor the shower-head holes are inclined at 30 deg to the spanwise direction, and are normal to the streamwise direction on the blade. Wilcox's k-omega turbulence model is used. The present study provides a much better comparison for the span-averaged heat transfer coefficient on the blade surface with the experimental data than an earlier analysis wherein coolant velocity and temperature distributions were specified at the hole exits rather than extending the computational domain into the hole-pipe and plenum. Details of the distributions of coolant velocity, temperature, k and omega at the hole exits are also presented.

Garg, Vijay K.↗
NASA NTRS1998

Heat Transfer on a Film-Cooled Rotating Blade Using a Two Equation Turbulence Model

A three-dimensional Navier-Stokes code has been used to compare the heat transfer coefficient on a film-cooled, rotating turbine blade. The blade chosen is the ACE rotor with five rows containing 93 film cooling holes covering the entire span. This is the only film-cooled rotating blade over which experimental data is available for comparison. Over 2.278 million grid points are used to compute the flow over the blade including the tip clearance region, using Coakley's q-omega turbulence model. Results are also compared with those obtained by Garg and Abhari (1997) using the zero-equation Baldwin-Lomax (B-L) model. A reasonably good comparison with the experimental data is obtained on the suction surface for both the turbulence models. At the leading edge, the B-L model yields a better comparison than the q-omega model. On the pressure surface, however, the comparison between the experimental data and the prediction from either turbulence model is poor. A potential reason for the discrepancy on the pressure surface could be the presence of unsteady effects due to stator-rotor interaction in the experiments which are not modeled in the present computations. Prediction using the two-equation model is in general poorer than that using the zero-equation model, while the former requires at least 40% more computational resources.

Garg, Vijay K.↗
NASA NTRS1997

Comparison of Two-Equation Turbulence Models for Prediction of Heat Transfer on Film-Cooled Turbine Blades

A three-dimensional Navier-Stokes code has been used to compute the heat transfer coefficient on two film-cooled turbine blades, namely, the VKI rotor with six rows of cooling holes, including three rows on the shower head and the C3X vane with nine rows of holes, including five rows on the shower head. Predictions of heat transfer coefficient at the blade surface using three two-equation turbulence model specifically, Coakley's q-omega model, Chien's k-epsilon model and Wilcox's k-omega model with Menter's modifications, have been compared with the experimental data of Camci and Arts for the VKI rotor, and of Hylton et al. for the C3X vane along with predictions using the Baldwin-Lomar (B-L) model taken from Garg and Gaugler. It is found that for the cases considered here the two equation models predict the blade heat transfer somewhat better than the B-L model except immediately downstream of the film-cooled holes on the suction surface of the VKI rotor, and over most of the suction surface of the C3X vane. However, all two-equation models require 40% more computer core than the B-L model for solution, and while the q-omega and k-epsilon models need 40% more computer time than the B-L model the k-omega model requires at least 65% more time because of the slower rate of convergence. It is found that the heat transfer coefficient exhibit a strong spanwise as well as streamwise variation for both blades and all turbulence models.

Garg, Vijay K.↗
NASA NTRS1999

Simultaneous Active Source Control of Blade Row Interaction Generated Discrete Tones

Unsteady blade row interactions in turbomachines generate discrete-frequency tones at blade pass frequency and its harmonics. Specific circumferential acoustic modes are generated. However, only certain of these modes propagate upstream and downstream to the far field, with these the discrete frequency noise received by an observer. This paper is directed at experimentally demonstrating the viability of active noise control utilizing active airfoils to generate propagating spatial modes that interact with and simultaneously cancel the upstream and downstream propagating acoustic modes. This is accomplished by means of fundamental experiments performed in the Purdue Annular Cascade Research Facility configured with 16 rotor blades and 18 stator vanes. At blade pass frequency, only the k(sub 0) = -2 spatial mode propagates. Significant simultaneous noise reductions are achieved for these upstream and downstream propagating spatial modes over a wide range of operating conditions.

Sawyer, Scott↗
NASA NTRS2000

Unsteady High Turbulence Effects on Turbine Blade Film Cooling Heat Transfer Performance Using a Transient Liquid Crystal Technique

Unsteady wake effect, with and without trailing edge ejection, on detailed heat transfer coefficient and film cooling effectiveness distributions is presented for a downstream film-cooled gas turbine blade. Tests were performed on a five-blade linear cascade at an exit Reynolds number of 5.3 x 10(exp 5). Upstream unsteady wakes were simulated using a spoke-wheel type wake generator. Coolant blowing ratio was varied from 0.4 to 1.2; air and CO2 were used as coolants to simulate different density ratios. Surface heat transfer and film effectiveness distributions were obtained using a transient liquid crystal technique; coolant temperature profiles were determined with a cold wire technique. Results show that Nusselt numbers for a film cooled blade are much higher compared to a blade without film injection. Unsteady wake slightly enhances Nusselt numbers but significantly reduces film effectiveness versus no wake cases. Nusselt numbers increase only slic,htly but film cooling, effectiveness increases significantly with increasing, blowing ratio. Higher density coolant (CO2) provides higher effectiveness at higher blowing ratios (M = 1.2) whereas lower density coolant (Air) provides higher 0 effectiveness at lower blowing ratios (M = 0.8). Trailing edge ejection generally has more effect on film effectiveness than on the heat transfer, typically reducing film effectiveness and enhancing heat transfer. Similar data is also presented for a film cooled cylindrical leading edge model.

Han, J. C.↗
NASA NTRS2000

Blading System and Method For Controlling Structural Vibrations

A new blading system for controlling the structural vibrations in axial-flow compressors, turbines, or fans, as in aircraft engines and like turbomachines including a stator disc and a rotor disc is presented. The rotor disc defines several radial hubs that retain the rotor blading systems. Each blading system includes a blade formed of an airfoil, and a root attachment which is dimensioned to fit within, and to engage a corresponding hub. Viscoelastic dampers are selectively applied to the outer surfaces of the root attachment on which compressive or shear forces are likely to develop, intermediate the root attachment and the hub, for compression therebetween upon rotation of the rotor disc, in order to dampen structural vibrations. One advantage presented by the viscoelastic dampers lies in its simplicity, efficiency, cost effectiveness, and its ability to be retrofitted into existing turbomachines with minor surface treatment of the root attachments. Furthermore, since the dampers are not exposed to the inflowing airstream, they do not affect the aerodynamic performance of the turbomachine. Another feature of the damping system is that it provides a significant source of damping to minimize destructive structural vibrations, thereby increasing the durability of the turbomachine, and reducing acoustic noise accompanying high amplitude vibrations.

Nguyen, Nhan↗
NASA NTRS1999

Heat Transfer on a Film-Cooled Rotating Blade Using Different Turbulence Models

A three-dimensional Navier Stokes code has been used to compute the heat transfer coefficient on a film-cooled, rotating turbine blade. The blade chosen is the ACE rotor with five rows containing 93 film cooling holes covering the entire span. This is the only film-cooled rotating blade over which experimental data is available for comparison. Over 2.278 million grid points are used to compute the flow over the blade including the tip clearance region. using Wilcox's k-omega model, Coakley's q-omega model, and the zero-equation Baldwin-Lomax (B-L) model. A reasonably good comparison with the experimental data is obtained on the suction surface for all the turbulence models. At the leading edge, the B-L model yields a better comparison than tile two-equation models. On the pressure surface however the comparison between the experimental data and the prediction from the k-omega model is much better than from the other two models. Overall, the k-omega model provides the best comparison with the experimental data. However, the two-equation models require at least 40% more computational resources than the B-L model.

Garg, Vijay K.↗
NASA NTRS2001

Two-Equation Turbulence Models for Prediction of Heat Transfer on a Transonic Turbine Blade

Two versions of the two-equation k-omega model and a shear stress transport (SST) model are used in a three-dimensional, multi-block, Navier-Stokes code to compare the detailed heat transfer measurements on a transonic turbine blade. It is found that the SST model resolves the passage vortex better on the suction side of the blade, thus yielding a better comparison with the experimental data than either of the k-w models. However, the comparison is still deficient on the suction side of the blade. Use of the SST model does require the computation of distance from a wall, which for a multiblock grid, such as in the present case, can be complicated. However, a relatively easy fix for this problem was devised. Also addressed are issues such as (1) computation of the production term in the turbulence equations for aerodynamic applications, and (2) the relation between the computational and experimental values for the turbulence length scale, and its influence on the passage vortex on the suction side of the turbine blade.

Garg, Vijay K.↗
NASA NTRS2001

Heat Transfer and Flow on the Blade Tip of a Gas Turbine Equipped with a Mean-Camberline Strip

Experimental and computational studies have 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 (greater than 100 MW) In a previous work the numerical heat transfer results for a sharp edge blade tip and a radiused blade tip were presented. More recently several other tip treatments have been considered for which the tip heat transfer has been measured and documented. This paper is concerned with the numerical prediction of the tip surface heat transfer for radiused blade tip equipped with mean-camberline strip (or "squealer" as it is often called). The heat transfer results are compared with the experimental results and discussed. The effectiveness of the mean-camberline strip in reducing the tip leakage and the tip heat transfer as compared to a radiused edge tip and sharp edge tip was studied. The calculations show that the sharp edge tip works best (among the cases considered) in reducing the tip leakage flow and the tip heat transfer.

Ameri, A.A.↗
NASA NTRS2001

Prediction of Relaminarization Effects on Turbine Blade Heat Transfer

An approach to predicting turbine blade heat transfer when turbulent flow relaminarizes due to strong favorable pressure gradients is described. Relaminarization is more likely to occur on the pressure side of a rotor blade. While stators also have strong favorable pressure gradients, the pressure surface is less likely to become turbulent at low to moderate Reynolds numbers. Accounting for the effects of relaminarization for blade heat transfer can substantially reduce the predicted rotor surface heat transfer. This in turn can lead to reduced rotor cooling requirements. Two-dimensional midspan Navier-Stokes analyses were done for each of eighteen test cases using eleven different turbulence models. Results showed that including relaminarization effects generally improved the agreement with experimental data. The results of this work indicate that relatively small changes in rotor shape can be utilized to extend the likelihood of relaminarization to high Reynolds numbers. Predictions showing how rotor blade heat transfer at a high Reynolds number can be reduced through relaminarization are given.

Boyle, R. J.↗
NASA NTRS2002

Rocket Engine Turbine Blade Surface Pressure Distributions Experiment and Computations

Understanding the unsteady aspects of turbine rotor flow fields is critical to successful future turbine designs. A technology program was conducted at NASA's Marshall Space Flight Center to increase the understanding of unsteady environments for rocket engine turbines. The experimental program involved instrumenting turbine rotor blades with miniature surface mounted high frequency response pressure transducers. The turbine model was then tested to measure the unsteady pressures on the rotor blades. The data obtained from the experimental program is unique in two respects. First, much more unsteady data was obtained (several minutes per set point) than has been possible in the past. Also, an extensive steady performance database existed for the turbine model. This allowed an evaluation of the effect of the on-blade instrumentation on the turbine's performance. A three-dimensional unsteady Navier-Stokes analysis was also used to blindly predict the unsteady flow field in the turbine at the design operating conditions and at +15 degrees relative incidence to the first-stage rotor. The predicted time-averaged and unsteady pressure distributions show good agreement with the experimental data. This unique data set, the lessons learned for acquiring this type of data, and the improvements made to the data analysis and prediction tools are contributing significantly to current Space Launch Initiative turbine airflow test and blade surface pressure prediction efforts.

Hudson, Susan T.↗