Supersonic Through-flow Fan Research Articles

Design optimization of a supersonic through-flow fan rotor based on the blade profiles

Abstract In this study, the authors propose an optimization process to design the baseline rotor of a supersonic through-flow fan (STFF) at an inlet Mach number of 2.0 based on Genetic Algorithm. Unlike the improvement in performance brought about by the pre-compression of conventional supersonic profiles in the presence of axial pressure flow, pre-compression did not help improve the performance of the rotor of the STFF. The efficiency of elements of the blade at spanwise heights of 10 %, 50 %, and 90 % increased by 2.47 %, 1.95 %, and 2.49 %, respectively. The performance of the rotor of the STFF that was reconstructed by stacking the optimized elements of the blade was improved at the design point as well as in off-design conditions by using three-dimensional computational fluid dynamics (CFD) simulations. The performance of the blade also improved considerably, with increases of by 2.46 % and 9.59 % in its isentropic efficiency and the overall pressure ratio, respectively.

Active subspace-based performance analysis of supersonic through-flow fan rotor

Abstract This article delves into the intricate relationship between the modeling parameters of an axial supersonic through-flow fan (STFF) rotor and its performance based on an active subspace method. It considers the influence of the STFF rotor key parameters on its performance. Implementing the active subspace method generates cloud maps to visualize the performance of the STFF rotor. Moreover, this study investigates the correlation between Bezier curve variables for constructing blade angles and critical performance metrics, such as the total pressure ratio and isentropic efficiency. After a 50 % chord length, the Bezier control point parameters dominate the effect on the total pressure ratio with a linear relationship. This article provides comparative flow field analyses on the blade elements with different performances under three working conditions. Under the constraint of the blade chamber turning angle, there is a linear relationship between the upper and lower limits of the isentropic efficiency distribution and total pressure ratio. This study shows that the total pressure ratio, installation angle, and maximum deflection value are positively associated. Further analysis provides an empirical formula.

Effect of inlet Mach number on performance and flow structure of an axial supersonic through-flow fan cascade

Ensuring the high efficiency and stable operation of a supersonic through-flow fan (STFF) in a wide range of inlet Mach numbers is of vital importance. The influence of inlet Mach number (M) ranging from 0.3 to 2.36 on the aerodynamic performance and flow structure in the STFF cascade is studied. The results indicate that at design incidence, the transonic inflow condition has a greater loss, a lower static pressure ratio, and a larger flow turning than the non-transonic inflow condition. The evolution of shock structure with the increasing inlet Mach number is as follows: shock-free → passage quasi-normal shock → leading edge (LE) local normal shock + λ shock → LE normal shock + trailing edge (TE) fishtail shock → LE bow shock + TE fishtail shock + multiple reflected oblique shocks → multiple oblique shocks. Under subsonic inflow conditions, the blade loss is dominated by boundary layer friction loss and wake loss. Under transonic inflow conditions, when the inlet Mach number rises from 0.66 to 0.86, the rear branch of passage λ shock impinges on the thickened separation region of the adjacent blade. This increases the loss. When the inlet Mach number rises to 1.06, the λ shock evolves into fishtail shocks, and the LE shock becomes stronger, reducing shock loss and separation loss significantly. As the inlet Mach number goes up to 1.26, the LE normal shock evolves into a bow shock, inducing a sharp reduction in LE shock loss. Under the supersonic inflow conditions, when the inlet Mach number rises from 1.96 to 2.36, the loss variation is mainly stemmed from LE bluntness loss, accounting for 66% of overall loss change.

Influence of Solidity on Aerodynamic Performance of Axial Supersonic Through-Flow Fan Cascades

To determine the effects of solidity on its aerodynamic performance and key flow structures, this study reports numerical simulations of the cascades of an axial supersonic through-flow fan with an inlet Mach number of 2.36 over a 15 deg range of incidence. The results indicate that, as the solidity increased, the overall loss first dropped and then increased. At the design incidence, loss at optimal solidity was 78% lower than the maximum loss. The overall loss was heavily dependent on the solidity but insensitive to incidence, especially at low solidity. At a critical solidity (2.38), the shock at the leading edge on the pressure side impinged on the suction surface near the trailing edge. At the optimal solidity, the wake loss was minimized and the number of reflected shocks decreased. The wake loss can be reduced by causing the shock to impinge on the suction surface near the throat of the passage. Moreover, this shock loss can be controlled by weakening its intensity. While reducing the local separation induced by the shock/boundary-layer interaction is ineffective for simultaneously reducing the loss of viscosity, reducing the loss of friction dominated by the surface area of the blade per unit of mass flow is effective.

Impacts of Tandem Configurations on the Aerodynamic Performance of an Axial Supersonic Through-Flow Fan Cascade

Abstract In the current study, the tandem blade technology is applied to an supersonic throughflow fan (STFF) tandem cascade for the first time and a 2D STFF tandem cascade is preliminarily designed. Through the modification design of the tandem airfoils and their configuration (axial overlap (AO) and percent pitch (PP)), the coefficients of total pressure loss and loading are reduced by 4% and 8.58%, respectively. Furthermore, the impact of tandem configurations on the performance is parametrically investigated by numerical simulations. The results indicate that compared with AO, the performance under design incidence is more sensitive to PP except for the cases with PP exceeding a threshold value (1.15). PP dominates the loss and load by controlling the evolution of the FB wake and the shock structure of FB and the rear one of the tandem blade (RB), while AO mainly adjusts the entire shock system structure through the change of virtual shape, resulting in the variation in load distribution between FB and RB. It is worth noting that the overall loading and the total loss remain unchanged with increasing AO except for the tandem configurations (PP = 1.05, AO ≤ −0.01), which make the flow structure in the gap region undergo a fundamental change. With the optimal tandem configuration (PP = 1.05, AO = −0.01) and the modified tandem blades (the ratios of chord length and camber for FB over RB are 0.67 and 0.5, respectively), the total pressure loss coefficient is further reduced by 19.7% in comparison with the preliminary tandem design.

Unsteady Quasi-One-Dimensional Nonlinear Dynamic Model of Supersonic Through-Flow Fan Surge

This paper presents the results of a computational fluid dynamic model of a supersonic through-flow fan (STF) in supersonic surge. The phenomenon of surge is well known for subsonic turbomachinery. However, for supersonic turbomachinery like the STF, a special type of supersonic surge occurs in which a shock oscillates through the fan, alternating locations between upstream and downstream of the fan. It is also possible for a shock to remain in a fan stage. This analysis focuses on the development of supersonic surge resulting from overpressuring the fan. A model of the STF was constructed around its steady-state experimental performance map using an existing unsteady, quasi-one-dimensional, inviscid, compressible flow code. An exit boundary condition was specified for the overpressuring that forces the shock to move upstream and interfere with normal fan operation. The effects of that interference, including a shock oscillating across the fan, are discussed.

Loss Reduction Using Riblets on a Supersonic Through-Flow Fan Blade Cascade

An experimental and computational study to determine the effects of riblets on the performance of the Supersonic Throughflow Fan (STF) cascade blades was performed. The cascade was tested in the Virginia Tech intermittent wind tunnel facility, where the Mach and Reynolds (based on chord) numbers were 2.36 and 4.8×106, respectively. The riblet sheets were symmetric v-grooved type and were applied onto the blade surfaces. Three different riblet heights were tested: 0.023, 0.033, and 0.051 mm. Riblet testing was conducted at design incidence as well as at off-design conditions (incidence angles: +5, −10 deg). Loss coefficients were measured and compared with a control test case where an equivalent thickness of smooth material was applied to the blade. Results show that at the design incidence, the riblet sheet with a height of 0.033 mm provides the optimal benefit, with a reduction of 8.5% in loss coefficient compared to the control case. Smaller effects were measured at the off-design conditions. In addition to the experimental study, a numerical investigation of the riblet effect on the STF cascade was conducted at design incidence. A simple method was developed to model riblet effects due to decrease in turbulent viscous drag and the delay of turbulent transition on the blades. Conclusions from numerical study indicate the 2/3 of the total decrease in losses are the result of delaying the transition location. The final 1/3 decrease in loss coefficient comes from the decrease in turbulent viscous losses.

Supersonic Through-Flow Fan Blade Cascade Studies

An investigation has been performed of the flow in a supersonic through-flow fan blade cascade. The blade shapes are those of the baseline supersonic through-flow fan (STF). Measurements were made at an inlet Mach number of 2.36 over a 15 deg range of incidence. Flowfield wave patterns were recorded using spark shadowgraph photography and steady-state instrumentation was used to measure blade surface pressure distributions and downstream flowfield. From these measurements, the integrated loss coefficients are presented as a function of incidence angle along with analysis indicating the source of losses in the STF cascade. The results are compared with calculations made using a two-dimensional, cell-centered, finite-volume, Navier-Stokes code with upwind options. Good general agreement is found at design conditions, with lesser agreement at off-design conditions. Analysis of the leading edge shock shows that the leading edge radius is a major source of losses in STF blades. Losses from the leading edge bluntness are convected downstream into the blade wake, and are difficult to distinguish from viscous losses. Shock losses are estimated to account for 70% to 80% of the losses in the STF cascade.

Use of pressure-sensitive paint for diagnostics in turbomachinery flows with shocks

The technology of pressure-sensitive paint (PSP) is well established in external aerodynamics. In internal flows in narrow channels and in turbomachinery cascades, however, there are still unresolved problems. In particular, the internal flows with complex shock structures inside highly curved channels present a challenge. It is not always easy and straightforward to distinguish between true signals and 'ghost' images due to multiple internal reflections in narrow channels. To address some of the problems, investigations were first carried out in a narrow supersonic channel of Mach number 2.5. A single wedge or a combination of two wedges was used to generate a complex shock wave structure in the flow. The experience gained in a small supersonic channel was used for surface pressure measurements on the stator vane of a supersonic throughflow fan. The experimental results for several fan operating conditions are shown in a concise form, including performance map test points, midspan static tap pressure distributions, and vane suction side pressure fields. Finally, the PSP technique was used in the NASA transonic flutter cascade to compliment flow visualization data and to acquire backwall pressure fields to assess the cascade flow periodicity. Lessons learned from this investigation and shortcomings of the PSP technology for internal flow application are presented in the conclusion of the paper.

Linear Calculation of Unsteady Aerodynamic Forces on Vibrating Cascade Blades Using TVD Scheme. 2nd Report. Application to Supersonic Through Flow Fan.

The TVD scheme combined with the finite difference method, which has been developed in the first part of the present papers to solve the linear unsteady flowfield of vibrating cascade blades, has been applied successfully to the vibrating supersonic through-flow fan cascade blades. Numerical example calculations indicate that the present calculation well predicts the general trend of dependences of the unsteady aerodynamic forces on various parameters, with little numerical oscillation in solutions. The steep changes in the flowfield accompanied by the spatial numerical oscillation predicted by the Lax Wendroff and the finite volume method are captured as less steep change in the present calculation.

Throughflow Method for Turbomachines Applicable for All Flow Regimes

A new axisymmetric throughflow method for analyzing and designing turbomachines is proposed. This method utilizes body-force terms to represent blade forces and viscous losses. The resulting equations of motion, which include these body-force terms, are cast in terms of conservative variables and are solved using a finite-volume time-stepping scheme. In the inverse mode, the swirl schedule in the bladed regions (i.e., the radius times the tangential velocity rVθ) is the primary specified flow quantity, and the corresponding blade shape is sought after. In the analysis mode, the blade geometry is specified and the flow solution is computed. The advantages of this throughflow method compared to the current family of streamline curvature and matrix methods are that the same code can be used for subsonic/transonic/supersonic throughflow velocities, and the proposed method has a shock capturing capability. This method is demonstrated for designing a supersonic throughflow fan stage and a transonic throughflow turbine stage.

Flow Field Analysis for the Supersonic Through-Flow Fan.

The implicit time-marching finite difference formulations of the Euler and Navier-Stokes equations using the upwind TVD scheme for two-dimensional compressible flows are applied to the flow field in supersonic through-flow fans. The results of the present Euler and Navier-Stokes equations are used to investigate the flow field for various geometrical and flow parameters. It is shown that the difference between the Euler calculation and the Navier-Stokes calculation is conspicuous in the region near shock waves or near the trailing-edge of the blades. Prediction of overall performance, however, is not susceptible to the two calculation models since the difference is limited to narrow region and is of small magnitude, unless the separation and reverse flow regions are generated.

Unsteady Lifting Surface Theory for Supersonic Through-Flow Fan.

The purpose of the present study is to predict analytically the unsteady loading on vibrating blades of a supersonic through-flow fan and investigate the aerodynamic instability. It is assumed that each blade operates with zero steady loading, and vibrates with small displacement amplitude, and a linearized unsteady lifting surface theory for a supersonic through-flow fan on the basis of the finite radial eigenfunction series approximation is developed. Numerical results of the unsteady loadings and unsteady aerodynamic works are presented. The bending vibration with zero mean loading is always stable, and the torsional instability heavily depends upon relative locations of the leading edge Mach wave reflection point and the torsion axis.

Unsteady Lifting Surface Theory Based on Double Linearization Concept for Supersonic Through-Flow Fan.

A full three-dimensional lifting surface theory based on the double linearization concept is developed for a rotating annular cascade model operating at axial supersonic velocity. It is assumed that each blade vibrates with infinitesimal displacement amplitude under small mean loading. Vibration modes normal and parallel to the blade chord are considered. Numerical results are presented to investigate the mean loading effects on the aerodynamic instability of the blade motion. It is found that the bending motion can be unstable due to the presence of mean loading. Some numerical results compared with strip theory predictions demonstrate significant three-dimensional effects on the unsteady aerodynamic force.

Performance Analysis of Supersonic Through-Flow Fan by the Lifting Surface Theory. 1st Report. Disturbance Flow Field and Determination of Blade Loadings.

The purpose of the present study, which is composed of two parts, is to predict analytically the steady performance of a supersonic through-flow fan in terms of design parameters. It is assumed that the fan cascade operates at supersnic axial velocity with a small angle of attack, camber and thickness. In 1st Report, the linearized lifting surface theory for a rotating annular cascade on the basis of the finite radial eigenfunction series approximation developed by the first author is extended to the supersonic through-flow fan model. Expressions of the disturbance flow field are derived, and the integral equation for the steady loading is numerically solved. The fundamental features of the blade surface pressure distribution dependent on the blade profile are demonstrated. The extensive parametric study of steady performance is presented in 2nd Report.

Unsteady Lifting Surface Theory for Supersonic Through-Flow Fan.

The purpose of this work is to predict analytically the unsteady loading on vibrating blades of a supersonic through-flow fan and investigate the aerodynamic instability. In order to deal with this problem, a linearized unsteady lifting surface theory on the basis of the finite radial eigenfunction series approxmation is developed. It is assumed that each blade operates with zero mean loading, and vibrates with a small displacement amplitude. Numerical results for pure bending, pure torsional and combined bending and torsional vibrations are presented to demonstrate influences of interblade phase angle, location of torsion axis, axial Mach number and reduced frequency on the flutter boundary.

Performance Analysis of Supersonic Through-Flow Fan by the Lifting Surface Theory. 2nd Report. Parametric Study.

Numerical solutions for steady loadings are presented for various combinations of prescribed design parameters. At low axial Mach numbers, high solidity and high camber with maximum camber position nearer to the trailing edge are desirable to attain higher total enthalpy increase. The contribution of the camber decreases with increasing axial Mach number. At high axial Mach numbers, the contribution of the angle of attack is predominant.

Influence of thickness and camber on the aeroelastic stability of supersonic throughflow fans

An engineering approach was used to include the nonlinear effects of thickness and camber in an analytical aeroelastic analysis of casecades in supersonic acial flow (supersonic leading-edge locus). A hybrid code using Lighthill's nonlinear piston theory and Lane's linear potential theory was developed to include these nonlinear effects. Lighthill's theory was used to calculate the unsteady pressures on the noninterference surface regions of the airfoils in cascade. Lane's theory was used to calculate the unsteady pressures on the remaining interference surface regions. Two airfoil profiles were investigated (a supersonic throughflow fan design and a NACA 66-206 airfoil with a sharp leading edge). Results show that compared with predictions of Lane's potential theory for flat plates, the inclusion of thickness (with or without camber) may increase or decrease the aeroelastic stability, depending on the airfoil geometry and operating conditions. When thickness effects are included in the aeroelastic analysis, inclusion of camber will influence the predicted stability in proportion to the magnitude of the added camber. The critical interblade phase angle, depending on the airfoil profile and operating conditions, may also be influenced by thickness and camber. Compared with predictions of Lane's linear potential theory, the inclusion of thickness and camber decreased the aerodynamic stiffness and increased the aerodynamic damping at Mach 2 and 2.95 for a cascade of supersonic throughflow fan airfoils oscillating 180 degrees out of phase at a reduced frequency of 0.1.

Application of Advanced Computational Codes in the Design of an Experiment for a Supersonic Throughflow Fan Rotor

Increased emphasis on sustained supersonic or hypersonic cruise has revived interest in the supersonic throughflow fan as a possible component in advanced propulsion systems. Use of a fan that can operate with a supersonic inlet axial Mach number is attractive from the standpoint of reducing the inlet losses incurred in diffusing the flow from a supersonic flight Mach number to a subsonic one at the fan face. The data base for components of this type is practically nonexistent; therefore, in order to furnish the required information for assessment of this type fan, a program has been initiated at the NASA Lewis Research Center to design, build, and test a fan rotor that operates with supersonic axial velocities from inlet to exit. This paper describes the design of the experiment using advanced computational codes to calculate the unique components required. The fan rotor has constant hub and tip radii and was designed for a pressure ratio of 2.7 with a tip speed of 457 m/s. The rotor was designed using existing turbomachinery design and analysis codes modified to handle fully supersonic axial flow through the rotor. A two-dimensional axisymmetric throughflow design code plus a blade element code were used to generate fan rotor velocity diagrams and blade shapes. A quasi-three-dimensional, thin shear layer Navier–Stokes code was used to assess the performance of the fan rotor blade shapes. The final design was stacked and checked for three-dimensional effects using a three-dimensional Euler code interactively coupled with a two-dimensional boundary layer code. A translating nozzle was designed to produce a uniform flow parallel to the fan up to the design axial Mach number of 2.0. The nozzle was designed with the three-dimensional Euler/interactive boundary layer code. The nozzle design in the expansion region was analyzed with a three-dimensional parabolized viscous code, which corroborated the results from the Euler code. A translating supersonic diffuser was designed using these same codes.