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.

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.

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 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.

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.

Strip distortion generator for simulating inlet flow distortion in gas turbine engine ground test facilities

A methodology has been developed to generate a non-uniform/distoited inlet flow field to test a gas turbine engine in ground test facilities.The distorted flow field is generated by positioning radial and circumferential strips of varying widths upstream of the Aerodynamic Interface Plane.The interacting wakes from these strips are used to generate a given target flow field.The approximate superposition of these wakes is investigated and used to construct the strip arrangement which is subsequently validated by computing the flow field by solving the Navier-Stokes equations.The strip geometry designed using the present methodology is able to produce the target Mach number distribution with a root-mean-square error of 5.06%.

Non-axisymmetric Flow Characteristics in Centrifugal Compressor

The flow field distribution in centrifugal compressor is significantly affected by the non-axisymmetric geometry structure of the volute.The experimental and numerical simulation methods were adopted in this work to study the compressor flow field distribution with different flow conditions.The results show that the pressure distribution in volute is characterized by the circumferential non-uniform phenomenon and the pressure fluctuation on the high static pressure zone propagates reversely to upstream,which results in the non-axisymmetric flow inside the compressor.The non-uniform level of pressure distribution in large flow condition is higher than that in small flow condition,its effect on the upstream flow field is also stronger.Additionally,the non-uniform circumferential pressure distribution in volute brings the non-axisymmetric flow at impeller outlet.In different flow conditions,the circumferential variation of the absolute flow angle at impeller outlet is also different.Meanwhile,the non-axisymmetric flow characteristics in internal impeller can be also reflected by the distribution of the mass flow.The high static pressure region of the volute corresponds to the decrease of mass flow in upstream blade channel,while the low static pressure zone of the volute corresponds to the increase of the mass flow.In small flow condition,the mass flow difference in the blade channel is bigger than that in the large flow condition.

Active flow control of S-duct by plasma synthetic jet

Flow separation and secondary flow in the S-duct of an aircraft engine cause severe pressure loss and airflow distortion at the outlet,lowering engine performance.Herein,a serial two-electrode plasma synthetic jet(PSJ)actuator array is used to actively control the flow field in the duct and improve its characteristics.The results show that the PSJ significantly increases the wall pressure recovery coefficient,suppresses flow separation,and improves the outlet pressure distortion.The primary and secondary orders of the influencing factors are as follows:control position>jet momentum coefficient>excitation frequency>jet configuration.The best jet control position is near the separation location,and the best jet configuration is the‘Λ’configuration.The higher the jet momentum coefficient and excitation frequency,the better the flow control.The wall pressure coefficient increases by up to 127.8%,and the outlet steady pressure distortion index decreases by 9.15%.The control mechanism is the direct energy injection into the flow boundary layer through a high-speed jet and the indirect control effect of the induced streamwise vortex.On the one hand,the PSJ suppresses flow separation by improving the ability of the boundary layer to resist the inverse pressure gradient.On the other hand,it reduces pressure distortion by decreasing the intensity of the secondary flow and weakening the backflow.This study thus provides a new technology for the active control of the flow-field characteristics in an S-duct and has significance for guiding the application of synthetic jet technology in S-ducts.