Exit Total Pressure Research Articles

Multi-objective optimization design of aerothermal performance for turbine blade squealer tip with inclined suction-side rim

To optimize the inclined rim design, multi-objective optimization was numerically conducted on the suction-side rim of the squealer tip in the GE-E3 rotor blade. The RANS equations, incorporating with the standard k-ω model, were numerically solved. Through the implementation of a Kriging surrogate model coupled with the NSGA-II evolutionary optimization technique, six geometric parameters on the suction-side rim went through iterative optimization. This process yielded three optimized squealer tip geometries on the Pareto front: Structure 1 with minimum total pressure loss, Structure 2 with optimal comprehensive aerothermal performance, and Structure 3 with minimum tip heat load. Comparative analysis revealed that with the inner and external wall inclined toward the cavity and passage respectively, the optimization results in a 3 %–11 % reduction in tip heat load, accompanied by an 8.5 %–10.2 % reduction in aerodynamic loss, facilitated by an attenuated upper passage vortex (UPV) and strengthened tip leakage vortex (TLV). The critical inclined region of the external rim wall, spanning 7 %–80 %, influences the exit total pressure loss distribution. The inclined inner wall upstream of 50 % Cax obstructs coolant jets from exiting the cavity, reducing the heat load on the cavity bottom. However, the inclined external wall from 66 % to 90 % Cax amplifies the impingement of the scraping vortex (SV), increasing local heat transfer. The thermal behavior of Structure 1 is sensitive to changes in tip clearance, whereas Structure 2 and 3 show more robust advantages. As the clearance increased, the aerodynamic losses associated with the TLV rise for all three optimized tips, but losses from the UPV remain similar. This study provides a novel approach for designing efficient squealer tip geometries incorporating an inclined rim.

INVERTED BRAYTON CYCLE ENGINE OPTIMIZATION FOR HYPERSONIC FLIGHT

Optimization of inverted Brayton cycle engine is made by particle swarm optimization method in this study. The optimum specific thrust value is reached by staying within the optimization constraints. When the total temperature of the cooling section is examined, it is seen that a temperature above the freezing temperature of the air is obtained. A very high specific thrust value, 451 N.s/kg is obtained at the hypersonic flight Mach Number (5 Mach) as a result of optimization. Temperature decrease with increasing altitude effect performance dramatically as positive. Low preburner exit total temperature and turbine total pressure ratio values and high afterburner exit total temperature values are desirable in terms of performance but values of total temperature at cooling section effect performance conversely in terms of specific thrust and specific fuel consumption

Effects of Ozone Addition on Multi-Wave Modes of Hydrogen–Air Rotating Detonations

Ozone addition presents a promising approach for optimizing and regulating both combustion and ignition mechanisms. In Rotating Detonation Engines (RDEs), investigating the impact of ozone addition is particularly important due to the fact of their unique operating conditions and potential for improved efficiency. This study explores the influence of ozone concentration, total temperature, and equivalent ratio on the combustion characteristics of a hydrogen–air mixture infused with ozone. Utilizing the mixture as a propellant, the combustion chamber of a continuous rotating detonation engine is replicated through an array of injection ports, with numerical simulations conducted to analyze the detonation wave combustion mode. Our results show that an increase in total temperature leads to an increase in the number of detonation waves. Incorporating a minor quantity of ozone can facilitate the ignition process for the detonation wave. Increasing the ozone content can result in the conversion from a single-wave to dual-wave or multi-wave mode, providing a more stable combustion interface. A low ozone concentration acts as an auxiliary ignition agent and can significantly shorten the induction time. As the total temperature increases, the detonation propagation velocity and the peak heat release rate both decrease concurrently, which leads to a decline in the exit total pressure and an augmentation in the specific impulse. Employing ozone exerts a minimal impact on the detonation propagation and the overall propulsion performance. The requirement for ozone-assisted initiation differs noticeably between rich and lean combustion.

Startability of truncated hypersonic wavecatcher intake

This paper compares the self-starting ability of an unmodified modular Busemann intake (UBI), with a modified Busemann-based intake (MBI), where the modification is designed to promote intake starting. Both intakes are designed on the wavecatching principle to allow overboard spillage during the starting process. Both intakes are designed for a flight Mach number of 5.19. The theoretical exit Mach number of both intakes is 3.11. Unsteady CFD calculations show that the UBI does not start spontaneously at any flight Mach number. The MBI does start at Mach 3.1. For the MBI Busemann flow, the theoretical exit total pressure recovery is 0.937 with a startability index Si=(Ae−Ais)/(Akan−Ais) of 1.507. The modified internal contraction results in an exit flow with an elliptic oblique shock and some non-uniformity. The paper presents, a) how an intake is truncated, b) how streamline tracing is applied to design a wavecatcher flowpath, and c) how the strong shock principle is applied to design a startable intake.

A linear cascade tunnel for flow investigations of steam turbine rotor tip blades in subsonic nucleating flows

The paper presents details of a unique experimental facility along with necessary accessories and instrumentation for testing steam turbine cascade blades in wet and nucleating steam. A steam turbine rotor tip cascade is chosen for flow investigations. Cascade inlet flow measurements show uniform conditions with dry air and steam and dry air mixture of different ratios. Exit flow surveys indicate that excellent flow periodicity is obtained. Blade surface static pressure and exit total pressure distributions are also presented with dry air and with steam and dry air mixture of different ratios as the working medium at an exit Mach number of 0.52.

Aerodynamic performance of profiled endwalls with upstream slot purge flow in a linear turbine cascade having pressure side separation

In aeroengines, purge flow directly fed from the compressor (which bypasses the combustor) is introduced through the disk space between blade rows to prevent the hot ingress. Higher quantity of purge gas fed through the wheel space can provide additional thermal protection to the passage endwall and blade surfaces. However, the interaction of purge flow with the mainstream flow leads to higher secondary losses. Secondary losses inside a turbine blade passage can be reduced effectively by endwall contouring. This paper presents computational investigation on the influence of non-axisymmetric endwall contouring over endwall secondary flow modification in the presence of purge flow with the pressure side bubble (PSB). The experimental analysis was conducted for the base case without purge and base case with purge (BCP) configurations having flat endwalls. The total pressure loss coefficient and exit yaw angle deviation were measured with the help of a five-hole pressure probe. Static pressure distribution over the blade midspan was obtained by 16 channel Scanivalve. Aerodynamic performances of three different profiled endwalls are numerically analyzed and are compared against the BCP configuration. The effects of different contoured endwall geometries on endwall static pressure distribution and secondary kinetic energy were also discussed. Analysis shows that in the first contoured endwall configuration (EC1), the formation of stagnation zones at a contour valley close to the suction surface causes the exit total pressure loss coefficient to increase. The shifting of the contour valley near to the pressure surface (EC2 configuration) has resulted in local acceleration of the diverted pressure side leg of the horseshoe vortex over the hump toward the end of the passage. In the third configuration (EC3 configuration), reduced valley depth and optimum hump height have effectively redistributed the endwall pitchwise pressure gradient. The increased static pressure coefficient at the endwall near to the pressure surface has eliminated the PSB formation. In addition, computational results of unsteady Reynolds averaged Navier–Stokes simulations are obtained for analyzing transient behavior of PSB, with more emphasis on its migration on the pressure surface and transport across the blade passage. The additional work done by the mainstream fluid to transport the low momentum PSB fluid has caused higher aerodynamic penalty at the blade exit region. In this viewpoint, the implementation of contoured endwalls has shown beneficial effects by eliminating the PSB and related secondary vortices. At 27% of axial chord downstream of the blade trailing edge, a 4.1% reduction in the total pressure loss coefficient was achieved with endwall contouring.

Computational study of the effect of cavity geometry on the supersonic mixing and combustion of ethylene

In this numerical study, the supersonic combustion of ethylene in three model combustor configurations namely, baseline (no cavity), square cavity and inclined cavity are investigated. To this end, 3D, compressible, turbulent, non-reacting (with fuel injection) and reacting flow calculations using single step and 10-step chemical kinetics in conjunction with a one equation turbulence model have been carried out. In the mixing study, predictions of the flow features with fuel injection, fuel trajectories, contours of total pressure loss along the combustor are compared between the combustor models. In the combustion study, contours of heat release are compared across the combustor models. Overall performance metrics such as mixing efficiency, total pressure loss and combustion efficiency are also compared between the combustor models. The comparison of combustor top wall static pressure and exit total pressure predictions with experimental data reported in the literature are also presented and discussed. The results clearly show that the model combustor with a shallow and inclined aft wall cavity has the highest residence time and maximum heat release. In addition, the role of a smaller L/D ratio cavity is shown to be minimal on the predictions of residence time and the heat release.

Experimental Study of Suction Flow Control Effectiveness in a Serpentine Intake

The intakes of modern aircraft are subjected to ever-increasing demands in their performance. Particularly, they are expected to carry out diffusion with the highest isentropic efficiency while subjected to aggressive geometry requirements arising from stealth considerations. To avoid a penalty in engine performance, the flow through intake needs to be controlled using various methods of flow control. In this study, a serpentine intake is studied experimentally and its performance compared with and without boundary layer suction. The performance parameters used are nondimensional total pressure loss coefficient and standard total pressure distortion descriptors. The effect is observed on surface pressure distributions, and inferences are made regarding separation location and extent. A detailed measurement at the exit plane shows flow structures that draw attention to secondary flows within the duct. Suction is applied at three different locations, spanning different number of ports along each location, comprising of ten unique configurations. The mass flow rate of suction employed ranges from 1.1% to 6.7% of mass flow rate at the inlet of the intake. The effect is seen on exit total pressure recovery as well as circumferential and radial distortion parameters. This is examined in the context of the location of the suction ports and amount of suction mass flow, by the deviation in surface pressure distributions, as well as the separation characteristics from the baseline case. The results show that applying suction far upstream of the separation point together with a modest amount of suction downstream results in the best performance.

Effects of vortex generator jet on corner separation/stall in high-turning compressor cascade

The effects of the vortex generator jet(VGJ)attached at the endwall on the corner separation/stall control are investigated by numerical simulation in a high-turning linear compressor cascade. The results show that the corner separation could be reduced significantly, which results in a wider operation range as well as a more uniform exit flow angle and total pressure profile. At the near-stall operation point, the maximum relative reduction of the total pressure loss is up to 32.5%,, whereas the jet mass ratio is less than 0.4%,. Based on the analysis of the detailed flow structure, three principal effects of the VGJ on the endwall cross flow and corner separation are identified. One is to increase the tangential velocity component opposite to cross flow, thus inhibiting the endwall secondary flow near the jet exit. The second is to suppress the pitchwise extension of the passage vortex as an air fence. The third is to sweep the low energy fluids towards the mainstream on the up-washed side and to transport the mainstream fluids to the endwall to reenergize the boundary layer on the down-washed side.

Numerical analysis on effects of inlet pressure and temperature non-uniformities on aero-thermal performance of a HP turbine

Unsteady numerical studies were conducted on the effects of combustor exit total pressure non-uniformity and hot streak (HS) on the aero-thermal characteristics of a high-pressure (HP) turbine stage. Firstly, four total pressure profiles including uniform, turbulent boundary layer, realistic engine combustor exit and radial reverse total pressure (higher total pressure near the endwall) were applied to reveal the vortex structures and heat transfer characteristics of the turbine. Then the four total pressure profiles coupled with inlet HS were specified to study the HS migration mechanisms. The results indicated that the turbulent boundary layer total pressure intensifies the passage vortices (PVs) while the higher total pressure near the endwall induces counter rotating vortices (CRVs), in both the stator and rotor passages. The heat transfer on the vane and blade surfaces was affected by the vortices and the extent of the impact directly depends on the intensity of the vortices, which is determined by the gradient of the total pressure. The heat transfer coefficients (HTCs) on the vane and blade pressure surface are increased by 104% and 20–30%, respectively, due to the CRV effects. The increase of tip leakage flow for the higher total pressure near the tip enhances the tip pressure side HTC by 4% comparing with the uniform case. The PVs and CRVs from the stator passage persistently exist in the rotor passage near the suction side endwall while have negligible influence on the rotor heat transfer. The migration paths of the hot streak are significantly affected by the PVs and CRVs in the stator passage, and the rotor blade surface temperature is dominated by both of the redistribution of HS in the vane passage and the vortices in the rotor passage. The PVs extend the HS radial migration and increase the vane and blade surface temperature. While the CRVs circulate the hot fluid away from the vane surface and increase the HS attenuation, thus decrease the vane and blade surface temperature significantly. The total pressure non-uniformity also decreases the turbine adiabatic efficiency by 0.6% at uniform inlet temperature condition.

A Computational Fluid Dynamics Study of Circumferential Groove Casing Treatment in a Transonic Axial Compressor

Numerical investigations were conducted to predict the performance of a transonic axial compressor rotor with circumferential groove casing treatment. The Notre Dame Transonic Axial Compressor (ND-TAC) was simulated at Tsinghua University with an in-house computational fluid dynamics (CFD) code (NSAWET) for this work. Experimental data from the ND-TAC were used to define the geometry, boundary conditions, and data sampling method for the numerical simulation. These efforts, combined with several unique simulation approaches, such as nonmatched grid boundary technology to treat the periodic boundaries and interfaces between groove grids and the passage grid, resulted in good agreement between the numerical and experimental results for overall compressor performance and radial profiles of exit total pressure. Efforts were made to study blade level flow mechanisms to determine how the casing treatment impacts the compressor's stall margin and performance. The flow structures in the passage, the tip gap, and the grooves as well as their mutual interactions were plotted and analyzed. The flow and momentum transport across the tip gap in the smooth wall and the casing treatment configurations were quantitatively compared.

Effect of tip geometry and tip clearance on aerodynamic performance of a linear compressor cascade

The tip leakage flow between a blade and a casing wall has a strong impact on compressor pressure rise capability, efficiency, and stability. Consequently, there is a strong motivation to look for means to minimize its impact on performance. This paper presents the potential of passive tip leakage flow control to increase the aerodynamic performance of highly loaded compressor blades. Experimental investigations on a linear compressor cascade equipped with blade winglets mounted to the blade tips have been carried out. Results for a variation of the tip clearance and the winglet geometry are presented. Current results indicate that the use of proper tip winglets in a compressor cascade can positively affect the local aerodynamic field by weakening the tip leakage vortex. Results also show that the suction-side winglets are aerodynamically superior to the pressure-side or combined winglets. The suction-side winglets are capable of reducing the exit total pressure loss associated with the tip leakage flow and the passage secondary flow to a significant degree.

Analysis of Aerodynamic Performance in an Annular Compressor Bowed Cascade with Large Camber Angles

The effects of positively bowed blade on the aerodynamic performance of annular compressor cascades with large camber angle were experimentally investigated under different incidences. The distributions of the exit total pressure loss and secondary flow vectors of compressor cascades were analyzed. The static pressure was measured by tapping on the cascade surfaces, and the ink-trace flow visualizations were conducted. The results show that the value of the optimum bowed angle and optimum bowed height decrease because of the increased losses at the mid-span with the increase of the caber angle. The C-shape static pressure distribution along the radial direction exists on the suction surface of the straight cascade with large r camber angles. When bowed blade is applied, the larger bowed angle and larger bowed height will further enhance the accumulation of the low-energy fluid at the mid-span, thus deteriorate the flow behavior. Under <TEX>$60^{\circ}$</TEX> camber angle, flow behavior near the end-wall region of some bowed cascades even deteriorates instead of improving because the blockage of the separated flow near the mid-span keeps the low-energy fluid near the end-walls from moving towards the mid-span region, and as a result, a rapid augmentation of the total loss is easy to take place under large bowed angle. With the increase of camber angle, the choice range of bowed angle corresponding to the best performance in different incidences become narrower.

Numerical analysis of supersonic gas-dynamic characteristic in laser cutting

The influence of the processing parameters on the dynamic characteristic of supersonic impinging jet in laser cutting is studied numerically. The numerical modeling of a supersonic jet impinging on a plate with a hole is presented to analyze the gas jet–workpiece interaction. The model is able to make quantitative predictions of the effect of the standoff distance and exit Mach number on the mass flow rate and the axial thrust. The numerical results show that the suitable cutting range is slightly different for different exit Mach number, but the optimal cutting parameter for certain exit total pressure is nearly changeless. So the better cut quality and capacity can be obtained mainly by setting the suitable standoff distance for a certain nozzle pressure.

Aerodynamic performance of bowed compressor cascades with different camber angles

The effects of a positively bowed blade on the aerodynamic performance of annular compressor cascades with different camber angles were experimentally investigated. The distributions of the exit total pressure loss and secondary flow vectors of the compressor cascades were analyzed. The static pressure was measured by tapping on the cascade surfaces, and the ink-trace flow visualizations were conducted. The results show that the value of the optimum bowed angle and optimum bowed height decrease because of the increased losses at the midspan with the increase of the camber angle. The C-shape static pressure distribution along the radial direction exists on the suction surface of the straight cascade with larger camber angles. When bowed blade is applied, the larger bowed angle and larger bowed height will further enhance the accumulation of the low-energy fluid at the mid-span, thus causing the flow behavior to deteriorate. Under 60° camber angle, flow behavior near the end-wall region of some bowed cascades even deteriorates instead of improving because the blockage of the separated flow near the mid-span keeps the low-energy fluid near the endwalls from moving towards the mid-span region. As a result, a rapid augmentation of the total loss can easily take place under a large bowed angle.

System-Level Performance Estimation of a Pulse Detonation Based Hybrid Engine

A key application for a Pulse detonation engine concept is envisioned as a hybrid engine, which replaces the combustor in a conventional gas turbine with a pulse detonation combustor (PDC). A limit-cycle model, based on quasi-unsteady computational fluid dynamics simulations, was developed to estimate the performance of a pressure-rise PDC in a hybrid engine to power a subsonic engine core. The parametric space considered for simulations of the PDC operation includes the mechanical compression or the flight conditions that determine the inlet pressure and the inlet temperature conditions, fill fraction, and purge fraction. The PDC cycle process time scales, including the overall operating frequency, were determined via limit-cycle simulations. The methodology for the estimation of the performance of the PDC considers the unsteady effects of PDC operation. These metrics include a ratio of time-averaged exit total pressure to inlet total pressure and a ratio of mass-averaged exit total enthalpy to inlet total enthalpy. This information can be presented as a performance map for the PDC, which was then integrated into a system-level cycle analysis model, using GATECYCLE, to estimate the propulsive performance of the hybrid engine. Three different analyses were performed. The first was a validation of the model against published data for a specific impulse. The second examined the performance of a PDC versus a traditional Brayton cycle for a fixed combustor exit temperature; the results show an increased efficiency of the PDC relative to the Brayton cycle. The third analysis performed was a detailed parametric study of varying engine conditions to examine the performance of the hybrid engine. The analysis has shown that increasing the purge fraction, which can reduce the overall PDC exit temperature, can simultaneously provide small increases in the overall system efficiency.

Aerodynamic character of partial squealer tip arrangements in an axial flow turbine. Part II: Detailed numerical aerodynamic field visualisations via three dimensional viscous flow simulations around a partial squealer tip

The present investigation deals with the numerical visualisation of tip leakage flow from 'partial-length squealer rims' that are constructed separately on the suction side in a single stage turbine rig. After describing the static pressure field on the tip surface with a squealer rim, the vortical flow features around the squealer rim and inside the tip gap are visualised in vertical planes. The leakage flow paths under the control of squealer rims are described in various planes parallel to the blade tip platform on which the squealer rims are attached. Turbine exit total pressure as calculated from the numerical model is presented for the baseline tip and tips with squealer rims. It is clearly shown that a partial squealer rim arrangement can be extremely effective in weakening the tip leakage vortex in a turbine facility provided that the squealer rim height is specified correctly. The study shows that the chordwise length of partial squealer rims near the suction side corner does not significantly affect the desensitisation process. The aerodynamic benefit of using a partial squealer tip arrangement located near the suction side of the tip platform in a turbine stage is demonstrated.

Aerodynamics of Tip Leakage Flows Near Partial Squealer Rims in an Axial Flow Turbine Stage

This paper deals with an experimental investigation of aerodynamic characteristics of full and partial-length squealer rims in a turbine stage. Full and partial-length squealer rims are investigated separately on the pressure side and on the suction side in the “Axial Flow Turbine Research Facility” (AFTRF) of the Pennsylvania State University. The streamwise length of these “partial squealer tips” and their chordwise position are varied to find an optimal aerodynamic tip configuration. The optimal configuration in this cold turbine study is defined as the one that is minimizing the stage exit total pressure defect in the tip vortex dominated zone. A new “channel arrangement” diverting some of the leakage flow into the trailing edge zone is also studied. Current results indicate that the use of “partial squealer rims” in axial flow turbines can positively affect the local aerodynamic field by weakening the tip leakage vortex. Results also show that the suction side partial squealers are aerodynamically superior to the pressure side squealers and the channel arrangement. The suction side partial squealers are capable of reducing the stage exit total pressure defect associated with the tip leakage flow to a significant degree.

Detection of separation in S-duct diffusers using shear sensitive liquid crystals

In the present experimental investigation, shear sensitive liquid crystals have been successfully used to study the flow characteristics and detect separation in two-dimensional Sduct diffusers of different curvatures. Tapered-fin vortex generators in two different orientations were used to control flow separation that was observed on one of the curved walls of the diffuser. The results were verified by conventional oil flow visualization technique and excellent agreement was observed. In addition to visualization, detailed measurements that included wall static pressure, skin friction, diffuser exit total pressure and velocity distributions were taken in a uniform inlet flow with Reynolds number of 3.49 × 10$^5$. These results are presented here in terms of skin friction distribution, distortion and total pressure loss coefficients. The extent of the separation zone (in terms of intensity of red distribution) in the diffuser with and without vortex generators (in both configurations) compared well with the Preston tube measurements. The present study demonstrates that shear sensitive liquid crystals can be efficiently used to study the flow physics in complex internal flows. In addition, the results also indicate that shear sensitive liquid crystals can be effectively used not only as flow visualization tool but also to gain quantitative information about the flow field in internal flows.

터빈성능시험에서 측정변수의 불확도 파급에 관한 연구

본 연구에서는 터빈 전효율의 불확도에 영향을 미치는 측정변수들의 불확도에 대한 영향을 실험의 결과로부터 분석하였다. 실험에 사용된 터빈은 축류형 3차원 단단터빈으로써 평균반경에서 반동도가 0.373이며 상온저압 상태에서 실험이 수행되었다. 입구전압력과 출구전압력에 의하여 발생되는 불확도는 전효율의 불확도에 전파되는 영향이 상당히 크게 나타나므로 상온저압시험에서는 높은 정밀도를 갖는 압력계를 사용하여야 한다. 측전변수의 불확도가 효율의 불확도에 차지하는 영향은 토오크에 의하여 발생되는 불확도가 가장 크게 나타났으며 회전수의 불확도가 가장 적었다. 토오크의 불확도 영향은 회전수의 증가에 따라 증가하는 경향이 있다. 전압력의 경우는 출그측에서의 불확도가 입구측에 비하여 영향력이 크게 나타났다. The effect of uncertainties caused by measured parameters, which are propagated to the uncertainty of total-to-total efficiency, are analyzed from a turbine performance test. The degree of reaction is 0.373 at the mean radius on a tested 3-D axial type turbine, and the performance test is conducted at the low pressure and cold temperature status. The uncertainty of turbine inlet and exit total pressure shows the strong propagation effect to the uncertainty of total-to-total efficiency. This means that a high precision pressure measuring system is required to reduce the uncertainty propagated by the pressure. In the uncertainty portion of each measured parameters to the uncertainty of total- to-total efficiency, the uncertainty by torque is the highest and the uncertainty by RPM is the lowest. In case of the total pressure, the effect of the uncertainty by torque is increased with the increasing RPM. The uncertainty of total pressure at the turbine exit is more important than that at the turbine exit.