Total Pressure Recovery Ratio Research Articles

Выбор параметров турбокомпрессора при проектировании ТРДД

The choice of the main parameters of a turbofan and turbocompressor at the initial stage of two-spool turbofan engine design is considered. To optimize rotation speeds, the number of stages, and determine the diametrical dimensions, line H is highlighted on Smith charts for compressors, and line L –for turbines. These lines are constructed through the tangent points of constant efficiency isolines and parabolas of the formТН/2ac=const and UL/2ac=const. Expressions are proposed that make it possible to determine the values of the total pressure recovery ratio σ at each point of the Smith diagrams for compressorand turbine stages (at corresponding temperature values). Taking this into account, an original method for determining the efficiency of compressors and turbines is proposed. It is shown how to use Smith diagrams to determine parameters not only for complete stages, but alsofor elementary stages in different sections along the height of the flow path. The proposed method makes it possible to maximize the efficiency of turbines due to the appropriate choice of the number of stages, the distribution of total head rise and the position of the mass-average points “mid” for the stages on the Smith diagrams. It is shown how these methods are used when fine-tuning compressors and turbines based on the results of turbofan engine tests

Design studies on NAGIC flowfield with application to integrated irregular-shaped supersonic inlet

Fighters have increasingly demanding performance on both aerodynamic and stealth. To adapt the design requirement of various irregular-shaped supersonic inlets integrated with flat fuselages, a design method for non-axisymmetric generalized internal conical (NAGIC) flowfield is discussed and then developed. By osculating axisymmetric flow concept and method of characteristics, the developed supersonic NAGIC flowfield can be inversely designed with arbitrary capture curves, controllable incident/terminal shock wave geometries, and post terminal shock wave pressure. Employing the streamline tracing technique, a set of external-compression inward-turning inlets with triangle-like flat capture shapes for supersonic stealth fighters are obtained to verify the design method. The study find that the internal contraction ratio and flow distortion of the inlet are reduced significantly by replacing the weak reflected shock wave in the original hypersonic NAGIC flowfield with a strong one. Thus, the compatibility between inlet and engine is improved in a relatively wide range. In addition, the irregular-shaped curved-compression inlet has high mass ratio and total pressure recovery ratio with negligible cross flow and uniform flow at throat. The performance of inlets with complex configurations can be further improved by matching the entrance geometries of basic flowfields and inlets.

ВЫБОР И ОПТИМИЗАЦИЯ ОСНОВНЫХ ПАРАМЕТРОВ ТУРБОКОМПРЕССОРОВ ПРИ ПРОЕКТИРОВАНИИ И ДОВОДКЕ ГТД

The choice and optimization of the rotational speed, the number of stages, the axial velocities distribution, reactivity, diametric dimensions and flow path shape, the blade rows width, changes in the height of the flow path in the compressors and turbines of the gas turbine engine are considered. The method for determining the total pressure recovery ratio and its relationship with the efficiency of blade rows, stages, compressors and turbines is proposed. It is shown that with the same value of this indicator in these elements, the maximum efficiency of compressors and turbines in the composition of turbocompressors is achieved. It is proposed to use this indicator at an early design stage to determine the optimal stages number and the work distribution between them, the change in work along the flow path height. It is shown how the position of the points for the stages and the optimal flow path shape are determined on the Smith diagrams. Various methods are proposed for parameters choosing of the blade profiles for the first, last and for intermediate steps. It is shown how these methods are used when fine-tuning compressors and turbines based on the results of GTE tests. The results of approbation of the developed design and refinement methods are given on examples of changing the geometry of the blade rows and the flow path in the replaceable flow path of the centrifugal blower for the R-16 "Ufa" gas compressor unit, the refinement of the 3-stage LPC for perspective two spool turbofan engine, comparison with the results of simulation in the DVIGw system , 2D and 3D CAD / CAE-modeling.

A proposed design method for supersonic inlet to improve performance parameters

A new structure for the compression surfaces of a supersonic inlet is proposed which has improved the target performance parameter i.e. total pressure recovery ratio. This idea resulted in development of a new type of supersonic inlet, utilized four ramps and a cone simultaneously as the compression surfaces. A prototype of the proposed inlet has been designed for a free stream Mach number of 3 and its performance has been evaluated via numerical simulation for both design and off-design conditions. The performance of the newly designed inlet has been compared to the existing experimental data of two double-cone inlets. The acquired data are further compared with analytical calculations of maximum reachable TPR for some conventional supersonic inlet types. The results confirmed the effectiveness of the main idea of the proposed design methodology. The total pressure recovery ratio of the newly designed inlet at its design condition, M∞=3, is calculated to be 76.8%.

Adjoint-Based Design Optimization of Vortex Generator in an S-Shaped Subsonic Inlet

This paper deals with an adjoint-based design optimization of vortex generators for the performance improvement of an S-shaped subsonic inlet, the Royal Aircraft Establishment intake model 2129. To enhance the flow quality entering the engine face maximally, the vortex generators are independently optimized with five design parameters per each vortex generator (a total of 55 design variables). To increase the design efficiency, the source term model of the vortex generator is employed. The original source term model, which does not reflect a small change in position and thus has difficulties in differentiation for sensitivity analysis, is modified into a differentiable source term model. To deal with a large number of design variables, the gradient-based design optimization method using the discrete adjoint approach is employed to minimize the distortion coefficient while maintaining the baseline total pressure recovery ratio. A total of five design cases are conducted to validate the proposed design approach, to obtain the optimized vortex generators, and to confirm their enhanced performance. Through the proposed design process,theperformanceofthetargetinletisremarkablyimproved,showingthatthedistortioncoefficientdecreases well over 70% while maintaining the total pressure recovery ratio.

초음속 유동 내 벤트 혼합기의 형상적 특성에 따른 성능 연구

본 논문은 새로운 개념의 초음속 혼합기인 벤트 혼합기의 형상적 특성에 따른 공력 특성을 연구하였다. 홀의 크기는 2 mm이며 혼합기 벽면에서 2 mm 떨어진 곳에 위치한 모델(case 1)과 혼합기 벽면 뒤쪽에 위치한 모델(case 2)의 경우 같은 전압력 회복율을 보였으며, 홀의 크기를 반으로 줄인 1 mm(case 3) 모델은 cases 1, 2에 비해 낮은 전압력 회복율을 보였다. 재순환 영역의 크기는 cases 1-3은 같지만 전단층 두께는 cases 1, 2가 case 3 보다 두꺼웠다. 재순환 영역 내 압력 손실의 경우 cases 1, 2은 case 3에 비해 낮은 압력 손실과 높은 속도 구배를 보였으며, 이는 재순환 영역 내 공기와 연료의 혼합을 증대시키는 요인이다. 재순환 영역 내로 유입 되는 유동에 의해 형성되는 박리 버블은 연소기의 전압력 회복율과 재순환 영역 내 압력 분포와 순환 유동에 영향을 미친다. 따라서 박리 버블 형성에 영향을 주는 유입 공기 유량이 벤트 혼합기 성능에 주요한 영향을 미치는 것을 알 수 있다. This paper studies the aerodynamic performance that the vent mixer-new conceptual supersonic mixer-showed with its geometric characteristics. The hole is 2 mm with 2 mm's distance from the wall in case 1 and with no distance in case 2. In case 3 die hole is 1 mm. Case 1 and case 2 showed the same total pressure recovery ratio, of which the case 3 was lower than that. While cases 1-3 had the same reattachment length, the shear layer was thicker in cases 1 and 2 than in case 3. Within the recirculation zone, cases 1 and 2 had lower pressure loss and higher velocity gradient difference than case 3-they enhance mixing between air and fuel. Separation bubble which is developed by the inflow into the recirculation zone has a significant effect on the total pressure recovery ratio in the combustor. Also separation bubble influences pressure distributions and recirculation flows in the recirculation zone. Therefore, inflow rate of air into the recirculation zone mainly affects the performance of vent mixer.

Hybrid Unstructured Navier-Stokes Simulation of Military Aircraft Configuration Including Inlet Airflow and Jet Exhaust Effects.

A CFD-based design system, CASPER (Computational Aerodynamics System for Performance Evaluation and Research), was developed at TRDI-JDA for advanced aircraft design at the next generation. In this paper, hybrid unstructured Reynolds-Averaged Navier-Stokes computations around a military aircraft, F-16 A, powered configurations including inlet airflow and jet exhaust effects of propulsion system were carried out as a practical example by using the CASPER. In order to validate the present CFD code, the present computed results are well compared to the wind-tunnel testing data with respect to total pressure recovery ratio on engine compressor face and aft-body pressure distributions at steady state. Thus, the present hybrid unstructured Navier-Stokes code is expected to be effective as one of design tools of advanced aircraft propulsion system.

極超音速飛しょう体の機体予圧縮に関する数値解析

Air intake is one of the most important components for an airbreathing propulsion system of supersonic and hypersonic vehicles. Air intake can be evaluated by air mass capture ratio and total pressure recovery ratio. In higher Mach number flight condition, larger total pressure losses occurs in the compression processes of air intake and reduces the propulsion performance. By utilizing the precompression coming from oblique shocks generated underneath vehicle forebody, a part of functions loaded in air intake can be substituted by the forebody precompression, thereby overall propulsive performance is able to be improved effectively. In the present paper, the precompression effects given by nose shape of forebody and geometrical arrangement of air intake underneath fuselage were analyzed by CFD calculation using 3-dimensional compressible Navier-Stokes equations.