Effects of total pressures and equivalence ratios on kerosene/air rotating detonation engines using a paralleling CE/SE method

In this paper,the kerosene/air rotating detonation engines(RDE)are numerically investigated,and the emphasis is laid on the effects of total pressures and equivalence ratios on the operation characteristics of RDE including the initiation,instabilities,and propulsive performance.A hybrid MPI t OpenMP parallel computing model is applied and it is proved to be able to obtain a more effective parallel performance on high performance computing(HPC)systems.A series of cases with the total pressure of 1 MPa,1.5 MPa,2 MPa,and the equivalence ratio of 0.9,1,1.4 are simulated.On one hand,the total pressure shows a significant impact on the instabilities of rotating detonation waves.The instability phenomenon is observed in cases with low total pressure(1 MPa)and weakened with the increase of the total pressure.The total pressure has a small impact on the detonation wave velocity and the specific impulse.On the other hand,the equivalence ratio shows a negligible influence on the instabilities,while it affects the ignition process and accounts for the detonation velocity deficit.It is more difficult to initiate rotating detonation waves directly in the lean fuel operation condition.Little difference was observed in the thrust with different equivalence ratios of 0.9,1,and 1.4.The highest specific impulse was obtained in the lean fuel cases,which is around 2700 s.The findings could provide insights into the understanding of the operation characteristics of kerosene/air RDE.

Numerical Research on Inlet Total Pressure Distortion in a Transonic Compressor with Non-Axisymmetric Stator Clearance

Inlet total pressure distortion has great adverse effects on the aero-engine performance. The distorted flow passes through the compressor and becomes non-uniform in the downstream blade rows. Different from previous studies based on the assumption of circumferential uniformity, this study aims to improve circumferential non-uniform flow with the non-axisymmetric structure. Non-axisymmetric stator clearance was adopted to resolve the effects of non-uniform flow caused by inlet total pressure distortion in this paper. The 9 stators with tip clearance were installed in the distorted region and the flow field structure and performance under different operating conditions was studied. The study finds that the non-axisymmetric compressor with 9 tip clearance stators can ensure compressor efficiency while improving compressor stability margin. What’s more, the separation range and strength in the distorted region can be reduced significantly and the anti-distortion capability of compressor can be enhanced.

船舶进气畸变模拟板结构设计与参数分析

Significant aerodynamic engine instability can occur during the operation of marine gas turbines as airflow enters the compressor through a 90°turning and causes inlet distortion.This study adopts the method of simulating board equivalence to provide the target distortion flow field for ship compressors.The characteristics of the flow field behind the simulated board are obtained through experiments and numerical simulations,through which the relationship between the height of the simulated board and the total pressure distortion is elucidated.Subsequently,the study summarizes the prediction formula to achieve a distortion prediction of 0.8%–7.8%.In addition,this work analyzes the effects of drilling methods and diameters on flow nonuniformity by drilling holes into the simulation board.The results indicate that drilling holes on the board can weaken the nonuniformity of the flow field within a certain range and change the distribution pattern of total pressure in the cross-section.Furthermore,the total pressure distortion no longer changes significantly when the number of holes is too large.The proposed double simulation board structure is capable of obtaining the following two types of distorted flow fields:symmetrical dual lowpressure zones and low-pressure zones with high distortion intensity at the compressor inlet.The distortion equivalent simulation method proposed in this work can obtain various types of distortion spectra,thereby meeting the distortion parameter requirements for the antidistortion testing of marine engines.

Modified Pressure Loss Model for T-junctions of Engine Exhaust Manifold

The T-junction model of engine exhaust manifolds significantly influences the simulation precision of the pressure wave and mass flow rate in the intake and exhaust manifolds of diesel engines. Current studies have focused on constant pressure models, constant static pressure models and pressure loss models. However, low model precision is a common disadvantage when simulating engine exhaust manifolds, particularly for turbocharged systems. To study the performance of junction flow, a cold wind tunnel experiment with high velocities at the junction of a diesel exhaust manifold is performed, and the variation in the pressure loss in the T-junction under different flow conditions is obtained. Despite the trend of the calculated total pressure loss coefficient, which is obtained by using the original pressure loss model and is the same as that obtained from the experimental results, large differences exist between the calculated and experimental values. Furthermore, the deviation becomes larger as the flow velocity increases. By improving the Vazsonyi formula considering the flow velocity and introducing the distribution function, a modified pressure loss model is established, which is suitable for a higher velocity range. Then, the new model is adopted to solve one-dimensional, unsteady flow in a D6114 turbocharged diesel engine. The calculated values are compared with the measured data, and the result shows that the simulation accuracy of the pressure wave before the turbine is improved by 4.3% with the modified pressure loss model because gas compressibility is considered when the flow velocities are high. The research results provide valuable information for further junction flow research, particularly the correction of the boundary condition in one-dimensional simulation models.

Aerodynamic Coupling Analysis of Counter⁃rotating Propfan and Inlet Flow Field

Taking a propfan engine as the research object,the CFD method was used for 3D modeling and unsteady slip flow for numerical calculation.The propfan rotation domain and the nacelle outside flow domain were meshed by using the partition splicing grid technology.Used the Reynolds⁃averaged of N⁃S equation,the Reynolds stress term uses the RNG turbulence model;and based on the slip grid method,numerical calculation of the flow field with different Mach numbers,front and rear blade angles and engine state were carried out;and the change law of propeller fan characteristics and the influence of slip flow on the inlet flow field were analyzed.The blade angle was the key parameter of the propeller fan characteristic conditions.When the blade angle increases from 41°to 50°,the thrust coefficient increases by 31.2%,and the power coefficient increases by 33.4%;in the climbing state of the propeller fan,the maximum total pressure distortion at the inlet port of 6.8%;the cross section is less affected by the slip flow of the propfan;and the pressure distribution is relatively uniform,but the area of the flow channel is small.The research results can provide a solution for the matching of the counter⁃rotating propeller fan and the engine and the arrangement of the air inlet measuring rake.

Tip Clearance Flows in Turbine Cascades

This article describes the effects of some factors on the tip clearance flow in axial linear turbine cascades. The measurements of the total pressure loss coefficient are made at the cascade outlets by using a five-hole probe at exit Mach numbers of 0.10, 0.14 and 0.19. At each exit Mach number, experiments are performed at the tip clearance heights of 1.0%, 1.5%, 2.0%, 2.5% and 3.0% of the blade height. The effects of the non-uniform tip clearance height of each blade in the pitchwise direction are also studied. The results show that at a given tip clearance height, generally, total pressure loss rises with exit Mach numbers proportionally. At a fixed exit Mach number, the total pressure loss augments nearly proportionally as the tip clearance height increases. The increased tip clearance heights in the tip regions of two adjacent blades are to be blame for the larger clearance loss of the center blade. Compared to the effects of the tip clearance height, the effects of the exit Mach number and the pitchwise variation of the tip clearance height on the cascade total pressure loss are so less significant to be omitted.

CHARACTERISTICS OF FLOW IN A PARTIALLY DYNAMIC HELICOPTER INLET

An experimental study of the flow in a helicopter inlet with front output shaft and partial flow dynamic head is conducted in low speed wind tunnel. The flow characters of the inlet in the range of the yaw angle from 0~135°are presented in this paper. The static pressure distributions along the duct, distortions of the flow field at the outlet section and total pressure recovery coefficients are measured and analyzed. The results show that this type of inlet has high total pressure recovery coefficients at a wide range of yaw angle. The regions of local flow separation and distortion are closely related to the yaw angle. It′s also found that the outlet section has the best characteristics at sideslip, and sharply deteriorated characteristics at the yawed flight with a yaw angle of more than 90°

总压畸变对涵道推力风扇气动性能影响的研究

Inlet distortion is one of the main factors for the degradation of aerodynamic performance and stability margin of the compressor in practical operation. Due to the change of the inlet shape and the large amount of inhalation of the body Boundary Layer, the ducted thrust fan of the Boundary Layer Ingestion (BLI) propulsion system inevitably works in the intake distortion condition. In this paper, the ducted thrust fan in a BLI propulsion system is taken as the research object. The influence of radial and circumferential total pressure distortion on the inlet section of the ducted thrust fan caused by boundary layer suction and inlet shape is studied by steady single channel and fullloop numerical simulation. The influence law of distortion intensity and distortion range of the two types of distortion patterns of the distortion map is analyzed emphatically. The results show that :(1) the greater the range and intensity of the radial total pressure distortion are, the more affected the performance of the ducted thrust fan is;(2) The aero-dynamic performance decline amplitude of the ducted thrust fan increases with the increase of the intensity of the circumferential total pressure distortion;The transmission law of the circumferential total pressure distortion intensity along the inlet and outlet of the fan is almost the same. Different working conditions have influence on the attenuation degree of the circumferential total pressure distortion in the ducted thrust fan, and the attenuation range of the circumferential total pressure distortion in the design working condition is the largest.

Serpentine Inlet Performance Enhancement Using Vortex Generator Based Flow Control

In order to provide the line-of-sight blockage of the engine face for an advanced Uninhabited Combat Air Vehicle（UCAV）, a highly curved serpentine inlet is proposed and experimentally studied. Based on the static pressure distribution measurement along the wall, the flow separation is found at the top wall of the second S duct for the baseline inlet design, which yields a high flow distortion at the exit plane. To improve the flow uniformity, a single array of vortex generators （VGs） is employed within the inlet. In this experimental study, the effects of mass flow ratio, free stream Mach number, angle of attack and yaw on the performance of a serpentine inlet instrumented with VGs are obtained. Results indicate： （1） Compared with the baseline serpentine design without flow control, the application of the VGs promotes the mixing of core flow and the low momentum flow in the boundary layer and thus prevents the flow separation. Under the design condition, the exit flow distortion （-↑△σ0） decreases from 11.7% to 2.3% by using the VGs. （2） With the descent of the free stream Mach number the total pressure loss decreases. However, the circular total pressure distortion increases. When the angle of attack rises from - 4° to 8°, the total pressure recovery and the circular total pressure distortion both go down. In addition, with the increase of yaw the total pressure recovery is fairly constant, while the circular total pressure distortion ascends gradually. （3） When Mao = 0.6-0.8, a = -4°-8° and β = 0°-6°, the total pressure recovery varies between 0.936 and 0. 961, the circular total pressure distortion coefficient varies between 1.4 % and 5.4 % and the synthesis distortion coefficient has a ranges from 3.8 % to 7.0 %. The experimental results confirm the excellent performance of the newly designed serpentine inlet incorporating VGs.

Numerical Analysis and Experimental Validation of a Submerged Inlet on the Plane Surface

In order to get a deep insight of a submerged inlet on the plane surface, the integrated flow field of the inlet and fuselage has been numerically studied. The investigation is mainly focused on the formation of the total pressure distribution at the exit of the inlet, the structure of the inner flow and the effects of the boundary layer along the fuselage on the performance of the inlet. Moreover, in comparison with the experimental data at different angles of attack, yaws and mass flow ratios, the reliabilities of the computational fluid dynamics（CFD） studied are verified. Results indicate： （1） the CFD results agree well with the experiment results and the relative errors of the total pressure coefficient is less than 1% ; （2） at the inlet＇s exit, the contour of total pressure obtained by CFD is similar to the experiment result except the contour in the low total pressure zone in CFD is slightly larger; （3） the secondary flow at the cross section behave as two counter-rotating vortices. Along the flow direction, the fields influenced by the vortex pair transport downstream and expand to the whole section at the exit; （4） the total pressure loss at the exit of the submerged inlet can be divided into external loss and internal loss. Usually, the external loss is greater than the internal loss, and both decrease with the augment of the Mach number at the exit. In addition, when the angle of attack ranges from -2° to 8°, the total pressure coefficient ascends gradually, due to the reduction of the external loss caused by the less boundary layer flow captured and the invisible change of the internal loss.

Numerical Investigation of Electrohydrodynamic (EHD) Flow Control in an S-Shaped Duct

An electrohydrodynamic （EHD） method, which is based on glow discharge plasma, is presented for flow control in an S-shaped duct. The research subject is an expanding channel with a constant width and a rectangular cross section. An equivalent divergence angle and basic function are introduced to build the three-dimensional model. Subsequently, the plasma physical models are simplified as the effects of electrical body force and work （done by the force） on the fluid near the wall. With the aid of FLUENT software, the source terms of momentum and energy are added to the Navier-Stokes equation. Finally, the original performance of three models （A, B and C） is studied, in which model A demonstrates better performance. Then EHD control based on model A is discussed. The results show that the EHD method is an effective way of reducing flow loss and improving uniformity at the duct exit. The innovation in this study is the assessment of the EHD control effect on the flow in an S-shaped duct. Both the parametric modeling of the S-shaped duct and the simplified models of plasma provide valuable information for future research on aircraft inlet ducts.

Numerical Investigation on the Secondary Flow Control by Using Splitters at Different Positions with Respect to the Main Blade

In turbomachinery,strong secondary flow can produce significant losses of total pressure near the endwall and reduce the efficiency of the considered turbomachine.In this study,splitters located at different positions with respect to the main blade have been used to reduce such losses and improve the efficiency of the outlet guide vane(OGV).Three different relative positions have been considered assuming a NACA 65-010 profile for both the main blade and the splitter.The numerical results indicate that splitters can effectively reduce the total pressure loss by suppressing the secondary flow around the main blade,but the splitters themselves also produce flow losses,which are caused by flow separation effects.

Numerical Simulation of Retrofit Scheme of a Boiler Tail Flue

Aiming at the problem that the total pressure loss of the flue of the electric precipitator of the 350 MW unit of a power plant to the inlet of the draft fan is too large,the numerical simulation software Fluent and the standard k-εmodel was used to simulate the flue,the results show that the main part of the flue mean total pressure loss is derived from the confluence header and elbow.In order to reduce the loss and consider the cost of transformation,the concept of twodimensional feature surface is established,gradually proposed three sets of flue transformation program and analysis of the flue transformation program drag reduction effect,the results show that the total reduction of the flue can be reduced from 486 Pa to 89 Pa and the reduction rate is 81.7%,which is the best solution;The concept of two-dimensional feature plane is helpful for quick condensing of flue;Double V-type structure of the convergence of the box has a better drag reduction effect.

Numerical Study on Aerodynamic Performance of CK Drone Aircraft Air Inlet in Maneuvering Flight

In view of the engineering background that CK drone aircraft needs modification and upgrading to improve its maneuvering performance,numerical research and analysis of air inlet aerodynamic performance are carried out.Firstly,based on the introduction of the theoretical knowledge involved in aircraft maneuvering flight,parameters such as aircraft attitude and engine mass flow etc.required for the aerodynamic performance calculation of CK drone aircraft air inlet are determined.By analyzing the test data of WP6 engine inlet distortion simulation board,the typical indexes are extracted as the basis for evaluating the air inlet performance of CK drone aircraft.Then,the aerodynamic characteristics of the inlet of CK drone aircraft under different maneuvering conditions are numerically studied,and the total pressure recovery coefficient and pressure distortion index of the outlet section are obtained.Several conclusions and suggestions are formed after the study.When CK drone aircraft flies at positive angle of attack,the inlet has good aerodynamic characteristics,which can meet the requirements of engine intake during high maneuverable flight.In the flight of negative angle of attack,the total pressure loss and pressure distortion at the outlet section of air inlet increase sharply,which cannot guarantee the stable working of the engine.On the premise that the aircraft attitude is satisfied,CK drone aircraft can use three engine thrust states of"Rated","Modified rated"and"Maximum"for high maneuverable flight.

Investigation of Changes in Flow Parameters along the Transport Line in Installations for Pneumatic Transport of Cotton

The article presents the results of a study of changes in the parameters of the flow, velocity and pressure of air along the cotton transportation line in pneumatic transport installations and their dependence on the parameters of the pipeline. The purpose of the research is to theoretically substantiate the choice of the pipeline diameter depending on the properties of the material (cotton), the required processing capacity and pipeline throughput. In the research, an analytical method was used to study changes in the pressure of the air flow along the transportation line at different pipeline diameters. It is established that, according to the existing calculation method, the flow pressure along the transportation line is reduced to zero. At the same time, various scientists have proposed various analytical and empirical dependencies, which, in turn, give different ideas about the resistance of the inner surface of the pipes to the movement of air and material flow and do not describe the real situation, especially when transporting material from a long distance. This requires in-depth theoretical and applied research to establish real patterns of changes in flow parameters along the material (cotton) transportation line.

Investigation of control effects of end-wall selfadaptive jet on three-dimensional corner separation of a highly loaded compressor cascade

To overcome the limitations posed by three-dimensional corner separation,this paper proposes a novel flow control technology known as passive End-Wall(EW)self-adaptive jet.Two single EW slotted schemes(EWS1 and EWS2),alongside a combined(COM)scheme featuring double EW slots,were investigated.The results reveal that the EW slot,driven by pressure differentials between the pressure and suction sides,can generate an adaptive jet with escalating velocity as the operational load increases.This high-speed jet effectively re-excites the local low-energy fluid,thereby mitigating the corner separation.Notably,the EWS1 slot,positioned near the blade leading edge,exhibits relatively low jet velocities at negative incidence angles,causing jet separation and exacerbating the corner separation.Besides,the EWS2 slot is close to the blade trailing edge,resulting in massive low-energy fluid accumulating and separating before the slot outlet at positive incidence angles.In contrast,the COM scheme emerges as the most effective solution for comprehensive corner separation control.It can significantly reduce the total pressure loss and improve the static pressure coefficient for the ORI blade at 0°-4° incidence angles,while causing minimal negative impact on the aerodynamic performance at negative incidence angles.Therefore,the corner stall is delayed,and the available incidence angle range is broadened from -10°--2°to -10°-4°.This holds substantial promise for advancing the aerodynamic performance,operational stability,and load capacity of future highly loaded compressors.

Experimental study on pressure and temperature distributions for low mass flux steam jet in subcooled water

A low mass flux steam jet in subcooled water was experimentally investigated.The transition of flow pattern from stable jet to condensation oscillation was observed at relatively high water temperature.The axial total pressures,the axial and radial temperature distributions were measured in the jet region.The results indicated that the pressure and temperature distributions were mainly influenced by the water temperature.The correlations corrected with water temperature were given to predict the dimen-sionless axial pressure peak distance and axial temperature distributions in the jet region,the results showed a good agreement between the predictions and experiments.Moreover,the self-similarity property of the radial temperature was obtained,which agreed well with Gauss distribution.In present work,all the dimensionless properties were mainly dependent on the water temperature but weakly on the nozzle size under a certain steam mass flux.

Design method for hypersonic bump inlet based on transverse pressure gradient

Transverse pressure gradient(TPG)is one of the key factors influencing the boundary layer airflow diversion in a bump inlet.This paper proposes a novel TPG-based hypersonic bump inlet design method.This method consists of two steps.First,a parametric optimization approach is employed to design a series of 2D inlets with various compression efficiencies.Then,according to the prescribed TPG,the optimized inlets are placed in different osculating planes to generate a 3D bump inlet.This method provides a means to directly control the aerodynamic parameters of the bump rather than the geometric parameters.By performing this method to a hypersonic chin inlet,a long and wide bump surface is formed in the compression wall,which leads to good integration of the bump/inlet.Results show that a part of the near-wall boundary layer flow is diverted by the bump,resulting in a slight decrease in the mass flow but a significant improvement in the total pressure recovery.In addition,the starting ability is significantly improved by adding the bump surface.Analysis reveals that the bump has a 3D rebuilding effect on the large-scale separation bubble of the unstarted inlet.Finally,a mass flow correction is performed on the designed bump inlet to increase the mass flow to full airflow capture.The results show that the mass flow rate of the corrected bump inlet reaches up to 0.9993,demonstrating that the correction method is effective.

On the Pressure Drop and the Velocity Distribution in the Cylindrical Vortex Chamber with Two Inlet Pipes for the Control of Vortex Flow

Vortex flow is applied to a cyclone dust collector, a vortex combustion chamber, and a vortex diode for vortex control. In order to apply the vortex flow to the industries, it is necessary to keep the stable flow condition and to estimate the response time of the transient flow process and also the intensity of the vortex flow. For control vortex flow, two types of vortex chamber with two inlet pipes were designed. One of them is to promote the vortex flow named as Co-Rotating Flow System and another one is to hinder the vortex flow named as Counter-Rotating Flow System. The pressure drops and the velocity distributions were measured for these vortex chambers. The estimation of the tangential velocity by the application of the angular momentum flux is compared with the measured velocity by a cylindrical Pitot-tube. The characteristics of the total pressure drop could be explained by introducing the circulation.

Design Optimization of Non-Axisymmetric Vane for an Axial Compressor under Inlet Distortion

The flow field at the inlet of compressors is generally encountered combined total pressure and swirl distortion for either aircraft engine with S-duct or gas turbine with lateral air intake.This inevitably deteriorates compressor aerodynamic performance,including not only the efficiency or pressure ratio but also the operation stability.In order to conquer this issue,appropriate measures such as integrating flow control techniques and modifying inlet or compressor design are of benefits.Due to this motivation,this article develops a full-annular two-dimensional(2D)and a partial-annular three-dimension(3D)optimization strategy for non-axisymmetric vane design.Firstly,two numerical simulation methods for evaluating performance of full-annular 2D vane and compressor with partial-annular 3D vane are developed.The swirl patterns at the inlet of a 1.5-stage axial compressor are analyzed and parametrized,and the parameterization is transferred to characterize the circumferential distribution of geometrical parameters of the vane profile.These approaches dramatically reduce computational simulation costs without violating the non-axisymmetric flow distortion patterns.Then various full-annular 2D sections at different radial locations are constructed as design space.The designed vane is reconstructed and 3D numerical simulations are performed to examine performance of the non-axisymmetric vane and the compressor with it.Also,partial annular 3D optimization is conducted for balancing compressor efficiency and stall margin.Results indicate that the designed non-axisymmetric vane based on full-annular optimization approach can decrease the vane total pressure loss under the considered inlet flow distortion,while those using partial-annular optimization achieve positive effects on compressor stall margin.