0

IN THIS VOLUME


Turbomachinery

1983;():V001T01A001. doi:10.1115/83-GT-2.
FREE TO VIEW

Two inverse design methods of a vaneless diffuser, with the inviscid core symmetry along radial mid-streamline, are presented. Inverse design models consist of an inviscid core and attached skewed turbulent boundary layers computed by Cebeci’s method. The displacement model was used for inverse design studies and compared with the entrainment model. Both inverse design models were corroborated with the published vaneless diffuser design featuring convergent contours. For the same inlet conditions the inverse design would provide attached flow while the tested diffuser had backflow region. Inverse design studies clarify the effect of the inlet swirl angle, Laval number, specific heat ratio, boundary layer blockage, skew angle, Reynolds number and skin friction coefficient on the diffuser contour generation for a radial pump or compressor stage. The inlet swirl angle was chosen to vary from 15–30°, the inlet Laval number from 0–1.1 and the inlet boundary layer blockage from 0.2% to 5% of inlet diffuser width.

Commentary by Dr. Valentin Fuster
1983;():V001T01A002. doi:10.1115/83-GT-3.
FREE TO VIEW

The pressure recovery of a low-solidity circular cascade diffuser of a centrifugal blower was predicted by a simple method combining a theory of circular cascade diffusers and that of vaneless diffusers and it was compared with a series of experiments. Furthermore the stall limit of the diffuser was studied. In order to improve the performance further, a series of tandem-cascade diffusers were tested. In these diffusers, the front row of the cascade was designed for a small flow rate while the rear row of the cascade was designed for a large flow rate so that the tandem cascade would accomplish good pressure recovery in a wide range of flow rate. Experimental results showed that the operating range was as wide as that of a vaneless diffuser and the pressure recovery was excellent at a small flow rate while it was somewhat better than that in a vaneless diffuser at a large flow rate.

Commentary by Dr. Valentin Fuster
1983;():V001T01A004. doi:10.1115/83-GT-8.
FREE TO VIEW

A new profile has recently been developed at Kraftwerk-Union (KWU), West Germany, to increase the efficiency of todays high pressure (HP) and intermediate pressure (IP) steam turbines. The profile was described using analytical functions. Therefore, the contour could be easily modified until the surface pressure distributions, which were calculated using a singularity method, fit a desired distribution. The experiments for the aerodynamic cascade data and blade surface pressure distributions were performed at the German Aerospace Research Establishment (DFVLR), Institute for Experimental Fluid Dynamics in Goettingen. Computer codes developed at the Institute enabled additional theoretical studies. The codes used in the study included a time-marching method for calculating the 2D-flow field and an integral method for calculating the boundary layer. Both theoretical and experimental results are discussed and compared with efficiency measurements derived from a multistage testrig at KWU.

Commentary by Dr. Valentin Fuster
1983;():V001T01A005. doi:10.1115/83-GT-9.
FREE TO VIEW

Results from an experimental study of the influence of the diffuser inlet shape on the performance of the diffuser and the whole compressor stage are presented. The investigations were carried out using a single stage centrifugal compressor. Three different vaned diffusers were tested. From detailed flow field measurements the influence of the diffuser inlet shape on the performance of the essential components of the compressor stage, i.e. the impeller, the diffuser, and the collecting chamber was analyzed. It is shown that the reaction of the vaned diffuser on the efficiency of the impeller is only weak but the losses in the collecting chamber are considerably affected by the used diffuser types.

Commentary by Dr. Valentin Fuster
1983;():V001T01A010. doi:10.1115/83-GT-22.
FREE TO VIEW

It has been observed that strong radially outward flow can be present on the pressure surface of an axial turbine rotor blade. This paper demonstrates that the relative eddy plays a major role in producing this radial flow. An analysis of the relative eddy indicates that it can explain observed trends both with blade incidence as well as with spanwise location on the blade. Suggestions are offered as to how the turbine designer might exercise some control over this aerodynamic mechanism.

Commentary by Dr. Valentin Fuster
1983;():V001T01A011. doi:10.1115/83-GT-23.
FREE TO VIEW

The operation of variable cycle gas turbines at negative incidence can result in highly three dimensional separated flows on the turbine rotor pressure surface. These flows can impact both performance and durability. The present program was conducted to experimentally study the behavior of surface flow on a large scale axial flow turbine rotor with incidence varying up to and including negative incidence separation. Fullspan pressure distributions and surface flow visualization were acquired over a range of incidence. The data indicate that at large negative incidence, pressure surface separation occurred and extended to 60 percent chord at midspan. These separated flows were simulated at midspan by applying potential flow theory to match the measured pressure distributions.

Commentary by Dr. Valentin Fuster
1983;():V001T01A012. doi:10.1115/83-GT-24.
FREE TO VIEW

A laser-Doppler anemometer was used to measure the three-dimensional velocity field within a typical turbine blade cascade. The blades had a 12.7 cm chord, a turning angle of 104.8°, and a shape conforming to the camber line of a commercial turboexpander. The cascade was operated at a Reynolds number of 1.25×105. Strong secondary velocities, ranging up to 35 percent of the primary flow velocity, were found, resulting from the development of counter-rotating vortices within the blade passages. Large midspan velocity defects in the primary flow were coincident with these high secondary flows. The secondary flow persisted throughout the near wake region.

Commentary by Dr. Valentin Fuster
1983;():V001T01A014. doi:10.1115/83-GT-34.
FREE TO VIEW

A unique non-rotating aerodynamic test facility is presented. It is designed for steady and unsteady transonic flow investigations in annular cascades. The fundamental advantage of this vehicle design, compared to a linear cascade facility, is the absence of boundary disturbances in the cascade and the self-adjustment of the flow periodicity. The downstream flow conditions can therefore be adjusted by the backpressure.

The nozzle consists of a radial-axial section, in which spiral axisymetric flow is generated. With the adjustment of the total pressure and swirl angles in two separate distributors, a continuous variation of Mach number between 0<M<1,4 and flow angles between 12°<β<70° (from tangential) can be obtained.

Further regulation of the flow quantities in the two separate distributors and boundary layer suction permits a continuous spanwise variation of the velocity and flow angle in the test section.

Commentary by Dr. Valentin Fuster
1983;():V001T01A015. doi:10.1115/83-GT-35.
FREE TO VIEW

Pressure recovery is examined for axisymmetric annular curved diffusers with various values of inlet/exit area ratio. The conditions of inlet flow include three swirl angles and two values of boundary layer blockage. Then, one of the diffuser is connected to two types of collectors with different size. The discharge pressure is decreased by the collector. The cause is not deterioration of diffuser performance due to asymmetric pressure distribution but formation of a cork screw type vortex flow inside the collector. An attempt to retard the vortex motion was successful by inserting obstructions in the collector. If a large size collector is used together with proper obstructions inside the collector, pressure-rise is possible in the collector rather than pressure drop.

Topics: Pressure , Diffusers
Commentary by Dr. Valentin Fuster
1983;():V001T01A016. doi:10.1115/83-GT-36.
FREE TO VIEW

An Euler code for the transonic flow through an axial flow rotor has been developed. The method of solution is an implicit time marching scheme and approximate factorization has been used to minimize the computations. Although the basic methods have been well publicized in the external aerodynamics literature, several modifications were found to be crucial in order to apply the methods to internal flows. the internal flow calculations appear to be much more susceptible to instabilities than the external flow calculations. A boundary-fitted coordinate system is used which is an adaptation of one due to Ives. The calculation of the metrics of the transformation proves to be extremely important and a revision of the numerical viscosity treatment enlarges and enhances the domain where converged solutions can be obtained. In particular, it was found that the metrics must be discretized in the same spatial fashion as the governing partial differential equation in order to avoid introducing source-like terms which would quickly destroy the solution. Results are presented for the two-dimensional case of flows through a cascade with inlet Mach numbers up to 0.76 and with outlet conditions prescribed and with a Kutta condition applied.

Topics: Rotors
Commentary by Dr. Valentin Fuster
1983;():V001T01A017. doi:10.1115/83-GT-37.
FREE TO VIEW

Aerodynamic tests on a novel concept of pre-combustor diffuser, particularly suited to can-annular configurations, are described. These tests show that the new arrangement avoids the unstable flow separations and some of the parasitic pressure losses which can be experienced in more conventional arrangements.

With the benefit of improved flow conditions, it is argued that the can-annular combustor arrangement has many advantages over other configurations, particularly for large engines, and where exhaust gas pollutants must be kept to a minimum.

Commentary by Dr. Valentin Fuster
1983;():V001T01A019. doi:10.1115/83-GT-41.
FREE TO VIEW

The flow field of a tilt-nacelle inlet-fan combination used for V/STOL aircraft is studied. Under certain flight conditions the inlet is subjected to high angles of attack and/or yaw. This produces a non-uniform or distorted flow field at the fan-face that can lead to large blade stresses. This paper presents an analytical approach to the coupled inlet-fan problem. The nacelle is modelled by a distribution of source panels and the fan by a distribution of radial vortices. A modified actuator disc with losses and a quasi-steady rotor response is used to derive the boundary condition at the fan-face. An example of the calculation is shown.

Commentary by Dr. Valentin Fuster
1983;():V001T01A020. doi:10.1115/83-GT-42.
FREE TO VIEW

Presented here is a simple method for designing optimum annular diffusers which enables the attainment of maximum static pressure recovery within a specified length. It is explained how the method can be adapted to account for compressible (but subsonic) flow.

Topics: Diffusers , Design
Commentary by Dr. Valentin Fuster
1983;():V001T01A021. doi:10.1115/83-GT-43.
FREE TO VIEW

Airflow tests have been conducted on an aerodynamic simulation of a combustor with pre-diffuser of compact configuration. The inlet Mach number throughout the tests was 0.35. The configuration was successful because of the attainment of a high pressure recovery, (Cp = 0.80), coupled with an exceptionally low total pressure loss (λ = 0.04). A useful analytical relationship is derived between the aerodynamic performance of combustor, compressor exit Mach number and diffuser performance.

Commentary by Dr. Valentin Fuster
1983;():V001T01A032. doi:10.1115/83-GT-70.
FREE TO VIEW

A procedure for obtaining entropy production rates from viscous flow calculations is described. The method is based on process thermodynamics; it allows loss production to be calculated in “irreversible equilibrium processes.”

The two-dimensional turbulent boundary layer of Samuel and Joubert is considered. Mean rates of entropy production are evaluated from measured data using rates of dissipation and rates of increase of turbulence kinetic energy. Calculations performed with the Moore Cascade Flow Program give good agreement with mean rates of entropy production and reveal details of the distribution of entropy production throughout the boundary layer. This method is used in Part II to reveal loss sources in calculations for a rectangular elbow.

Commentary by Dr. Valentin Fuster
1983;():V001T01A033. doi:10.1115/83-GT-71.
FREE TO VIEW

Entropy creation is studied in a three-dimensional turbulent flow in a rectangular elbow with 90 degrees of turning. In Part I, a method of evaluating entropy production rates in calculations of viscous flow was described. This method is used here to show where the calculated losses arise in the flow field.

The flow develops strong passage vortices as large shear layers at the inlet develop a streamwise component of vorticity. The entropy production in the shear layers is compared with the entropy creation in the wall boundary layers. The influence of the three-dimensional flow on the distributions of entropy production in the bend is discussed.

Commentary by Dr. Valentin Fuster
1983;():V001T01A034. doi:10.1115/83-GT-74.
FREE TO VIEW

The major problem for designing centrifugal compressors is to attain high stage efficiency as well as a wide operating range. High stage efficiency is customarily attained by the optimization of design parameters using a one-dimensional loss analysis including the relationship between the flow behavior and total pressure losses for limited types of compressors.

Commentary by Dr. Valentin Fuster
1983;():V001T01A035. doi:10.1115/83-GT-75.
FREE TO VIEW

The presence of solid particles in turbomachinery flow affects the component performance as well as its life. The subject of particulated flows can be broadly divided into three parts, namely, particle trajectories, the effect of particles on the aerodynamics of flow and material erosion. The first two aspects are investigated in this paper taking into account the viscosity of the carrier fluid.

The Lagrangian formulation is adopted for the particles, whereas the Eulerian approach is used for the continuous phase. The effect of particles is incorporated as interphase force terms in the fully incompressible stream function-vorticity form of the Navier-Stokes equations. The field analysis is based on the numerical integration of this equation over the rotor blade to blade stream channels. The numerical code used to solve the governing equations employs a nonorthogonal boundary fitted coordinate system that suits the most complicated blade geometries. The trajectories of the solid particles are determined including particle impacts with the blades. The particle rebounding velocity and direction after each impact is determined using semi-empirical correlations for the restitution ratios obtained experimentally. The method of analysis is applied to a radial inflow turbine. The effect of particles on the aerodynamics of the flow is studied by analyzing the fluid streamline pattern in the rotor blades with and without solid particles. The analysis is carried out for various particle concentrations.

Commentary by Dr. Valentin Fuster
1983;():V001T01A039. doi:10.1115/83-GT-92.
FREE TO VIEW

Analysis of three-dimensional, transonic, potential flow through a compressor rotor is investigated using the finite element method. The formulation of the equations and the generation of the corresponding finite element grid are presented. A constant-coefficient solution scheme is applied for the efficient solution of these equations. Numerical results are obtained for the flow through a 5.0” diameter turbocharger compressor wheel. Four different mass flow rates are considered including a transonic, near-choked flow case. Accuracy of the numerical results are tested by comparing with experimental measurements.

Commentary by Dr. Valentin Fuster
1983;():V001T01A040. doi:10.1115/83-GT-95.
FREE TO VIEW

A computational model for calculating the flow field, in the presence of solid particles, through a two dimensional compressor cascade is presented. The method treats the particle phase in the Lagrangian system and the fluid phase in the Eulerian system. The equations of momentum of the fluid phase are modified to account for the momentum exchange between the two phases. The resulting modified momentum and continuity equations are reduced to the conventional stream function-vorticity formulation.

The fluid phase is treated as incompressible and inviscid. Averaging of the particle phase properties, based on the trajectory analysis, over the control volume yields the fluid-particle momentum exchange at every grid point, where the flow field solution is desired. The inelastic rebound of the particles on impact of the airfoil surface is accounted for in the trajectory analysis.

The fluid phase stream function-vorticity equations are solved by a novel method. The computation is performed in a transformed plane where one of the transformed coordinates is the streamline itself. The transformed vorticity equation is solved by the space marching technique.

The analysis yields the change in the blade surface pressure distribution, the total pressure and velocity distributions. Based on this, it is possible to predict the decrease in compressor performance under the presence of particles.

Commentary by Dr. Valentin Fuster
1983;():V001T01A041. doi:10.1115/83-GT-116.
FREE TO VIEW

A conformal mapping technique, applicable to compressor and turbine cascades, is presented. The technique can be used to map the field surrounding the cascade into a rectangle, in which the image of the blade surface lies along one side of the rectangle, the image of the trailing edge is at two corners of the rectangle and the image of downstream infinity lies at the other two corners. These features make it possible to use the mapping in several contemporary flowfield-analysis codes, which carry out their calculations in such a rectangular plane. The mapping produces an O-type grid in the cascade plane.

Commentary by Dr. Valentin Fuster
1983;():V001T01A050. doi:10.1115/83-GT-134.
FREE TO VIEW

A computer code that generates shock-free transonic compressor cascade shapes while taking into account viscosity effects is developed. The mathematical model for the inviscid flow field is the full potential equation. The Kutta-Joukowski condition is satisfied by varying the free stream angle at downstream infinity. A boundary fitted computational grid of C-type is generated using a sequence of conformal mapping and nonorthogonal coordinate stretching and shearing transformations. The inviscid calculation is performed sequentially on up to four consecutively refined grids thereby accelerating the convergence of the solution process. The full potential equation is solved using a finite area technique and rotated, type-dependent finite differencing. Artificial viscosity of the first order is added in a fully conservative form. Shock-free cascade airfoil shapes are obtained using the fictitious gas concept of Sobieczky and the method of characteristics in the rheograph plane. Viscosity effects are incorporated via a boundary layer displacement thickness. The integral boundary layer code is based on Rotta’s turbulence model and assumes transition region of zero length.

Commentary by Dr. Valentin Fuster
1983;():V001T01A051. doi:10.1115/83-GT-135.
FREE TO VIEW

An analysis of internal potential flow is presented showing the existence of multiple potential solutions with shocks for a given mass flow rate. These solutions are related to non isentropic Euler solutions.

Inflow and outflow boundary conditions are proposed which uniquely determine the shock position allowing the calculation of potential flows which are either choked or have a supersonic inlet.

Numerical computations using a multigrid finite element approach are presented and compared with exact quasi-one-dimensional Euler solutions confirming the ability of the potential method to solve accurately supersonic and choked nozzle flows.

Commentary by Dr. Valentin Fuster
1983;():V001T01A053. doi:10.1115/83-GT-172.
FREE TO VIEW

Almost all process compressors manufactured today are built in accordance with the fixed geometry modular system.

It is not always possible to accurately predict the operational behaviour of a compressor particularly in such cases where a deviation from the standard solution is advisable or even necessary.

This paper describes and analyses by means of experimental measurements the influence of three different stage modifications (axial stage pitch, shortening of the vaneless annular diffusor, blading cut-back).

The results given here are directly related to engineering practice as they enable the design engineer to apply realistic efficiencies even for modified stages in the early planning stage.

Commentary by Dr. Valentin Fuster
1983;():V001T01A054. doi:10.1115/83-GT-204.
FREE TO VIEW

The clearance between a compressor or turbine case and rotating blades or shrouds has been measured in a variety of ways over the years with varying degrees of accuracy. However, one critical clearance which has defied active measurement heretofore has been that between the tip of a cantilevered stator vane and an adjacent drum rotor. A significant number of turbomachines under development, representing many millions of dollars, have met an inglorious end through unanticipated rubs which were experienced in this area.

The present paper describes how a device which was introduced by its manufacturer for conventional rotor tip or shroud clearance measurements was applied to measure stator-vane-to-drum clearance. The installation technique is explained, and the results of vane-to-drum measurements made with a research compressor are presented. Adaptations of the same technique should permit similar measurements to be made within a turbine, and possibly on rotating seals buried deep within a turbomachine.

Commentary by Dr. Valentin Fuster
1983;():V001T01A055. doi:10.1115/83-GT-208.
FREE TO VIEW

In order to verify a new controlled diffusion blade design concept, the stator of an existing transonic axial compressor stage was redesigned. Stator and equivalent cascade tests revealed the potential of such blades for a considerably higher aerodynamic loading than it has been applied up to now.

The design procedure is described and the results of plane cascade and stage testing are submitted including performance analysis of both, cascade and stator blade sections, at design and off-design operating conditions.

The blade profile shapes and cascade geometries were calculated by means of an inverse 2-dimensional method taking also into account the axial velocity density ratio (AVDR). This design concept is essentially based on prescribed blade pressure distributions which are optimized with respect to the blade boundary layer development.

The flow phenomena are illustrated by means of loss and flow turning investigations, blade pressure distributions and laser velocimetry data.

The test results reveal that the 2-D approach applied is quite promising for the 3-D stator blade design. Finally overall and blade element performance comparisons between the original NACA 65 profiled stator and the redesigned one demonstrate the favourable flow behaviour of the new stator as well as the great potential of the controlled diffusion blade concept.

Commentary by Dr. Valentin Fuster
1983;():V001T01A056. doi:10.1115/83-GT-209.
FREE TO VIEW

A transonic compressor rotor blade cascade was tested in order to elucidate the flow behaviour in the transonic regime and to determine the performance characteristic in the whole operating range of a rotor blade section. The experiments have been conducted in a transonic cascade wind tunnel, which enables tests even at sonic inlet velocities.

The influence of the upstream Mach number between 0.8 and 1.1 and the inlet flow angle between choking and stalling of the blade row was investigated. The effect of the axial velocity density ratio (AVDR) could be studied by applying an endwall suction device. Furthermore the level of the shock losses was determined from a wake analysis.

A final comparison of cascade losses and those of the corresponding rotor blade element shows good agreement which underlines the applicability of the cascade model in the design of axial flow turbomachines.

Commentary by Dr. Valentin Fuster
1983;():V001T01A057. doi:10.1115/83-GT-210.
FREE TO VIEW

The flow in a centrifugal impeller is analyzed by a quasi-three-dimensional streamline curvature method, by a three-dimensional Euler code, and by a three-dimensional finite element potential flow method. Comparison with Eckardt’s published data for a backswept impeller shows that full three-dimensional methods better predict the loading at the hub and shroud.

Commentary by Dr. Valentin Fuster
1983;():V001T01A058. doi:10.1115/83-GT-211.
FREE TO VIEW

A series of Controlled Diffusion Airfoils has been developed for multistage compressor application. These airfoils are designed analytically to be shock free at transonic Mach number and to avoid suction surface boundary layer separation for a range of inlet conditions necessary for stable compressor operation. They have demonstrated, in cascade testing, higher critical Mach number, higher incidence range, and higher loading capability than standard series airfoils designed for equivalent aerodynamic requirements. These airfoils have been shown, in single and multistage rig testing, to provide high efficiency, high loading capability, and ease of stage matching, leading to reduced development costs and improved surge margin. The Controlled Diffusion Airfoil profile shapes tend to have thicker leading and trailing edges than their standard series counterparts, leading to improved compressor durability.

Commentary by Dr. Valentin Fuster
1983;():V001T01A059. doi:10.1115/83-GT-215.
FREE TO VIEW

The flow from a high speed rotor in a rotor-first arrangement has been measured using a “dual-probe, digital sampling (DPDS)” technique. The flow field was found to be steady in rotor coordinates (periodic in machine coordinates) outside the rotor wake, and 3 components of velocity and the pressure field were determined in this area. The wake regions were unsteady. In these regions the measurements based on ensemble averages of multiple samples did not follow the behavior established during calibration in steady uniform freejet flow except near the wake center. The broadening of the wake and three dimensional effects in the flow field were measured at reduced throttle and increased speeds. The results serve to illustrate the potential of the measurement technique, on which the emphasis of the presentation is placed.

Commentary by Dr. Valentin Fuster
1983;():V001T01A060. doi:10.1115/83-GT-216.
FREE TO VIEW

A design trend evident in newly evolving aircraft turbine engines is a reduction in the aspect ratio of blading employed in fans, compressors, and turbines. As aspect ratio is reduced, various three-dimensional flow effects become significant which at higher aspect ratios could safely be neglected. This paper presents a new model for predicting the shock loss through a transonic or supersonic compressor blade row operating at peak efficiency. It differs from the classical Miller-Lewis-Hartmann normal shock model by taking into account the spanwise obliquity of the shock surface due to leading-edge sweep, blade twist, and solidity variation. The model is evaluated in combination with two test cases. Each was a low-aspect-ratio transonic stage which had exceeded its efficiency goals. Use of the revised shock loss model contributed 2.11 points to the efficiency of the first test case and 1.08 points to the efficiency of the second.

Commentary by Dr. Valentin Fuster
1983;():V001T01A061. doi:10.1115/83-GT-218.
FREE TO VIEW

A systemmatic series of tests has been conducted on a family of annular diffusers where the outer casing is maintained at constant diameter. Such a diffuser is typical of turbine exits. Data, in the form of static pressure recovery coefficient is plotted against diffuser length for several different designs of centerbody closure.

It has been shown that such diffusers can have short length centerbodies for which a set of design guides has been established.

Commentary by Dr. Valentin Fuster

Sorry! You do not have access to this content. For assistance or to subscribe, please contact us:

  • TELEPHONE: 1-800-843-2763 (Toll-free in the USA)
  • EMAIL: asmedigitalcollection@asme.org
Sign In