Influence of Propeller Slipstream on the Flow Field of S-Shaped Intake

For turboprop engine, the S-shaped intake affects the engine performance and the propeller is not far in front of the inlet of the S-shaped intake, so the slipstream inevitably affects the flow field in the S-shaped intake and the engine performance. Here, an S-shaped intake with/without propeller is studied by solving Reynolds-averaged Navier-Stokes equation employed SST k-ω turbulence model. The results are presented as time-averaged results and transient results. By comparing the flow field in S-shaped intake with/without propeller, the transient results show that total pressure recovery coefficient and distortion coefficient on the AIP section vary periodically with time. The time-averaged results show that the influence of propeller slipstream on the performance of S-shaped intake is mainly circumferential interference and streamwise interference. Circumferential interference mainly affects the secondary flow in the S-shaped intake and then affects the airflow uniformity; the streamwise interference mainly affects the streamwise flow separation in the S-shaped intake and then affects the total pressure recovery. The total pressure recovery coefficient on the AIP section for the S-shaped intake with propeller is 1%-2.5% higher than that for S-shaped intake without propeller, and the total pressure distortion coefficient on the AIP section for the S-shaped intake with propeller is 1%-12% higher than that for the S-shaped intake without propeller. However, compared with the free stream flow velocity ( Ma = 0.527 ), the influence of the propeller slipstream belongs to the category of small disturbance, which is acceptable for engineering applications.

Обґрунтування моделі турбулентної в’язкості для дослідження характеристик співвісного гвинтовентилятора і вхідного пристрою ГТД

The issues of substantiation of the most rational, based on adequacy, model of turbulent viscosity for mathematical modeling of the flow near the propfan and in the inlet of the turbine-propeller engine are considered. It was found that at present there is no universal turbulence model for determining the parameters of the boundary layer, energy loss in the flow, and laminar-turbulent transition. Analysis of the results of previous studies showed that there is a need to select and justify a turbulent viscosity model for each type of research object. The task of modeling the flow near the propfan and in the inlet device of the power plant was performed using the ANSYS CFX software product, which allows using various standard mathematical models and tools for modeling turbulent flow. The object of research is an annular axial inlet device, in front of which there is a coaxial propfan with two rows of propellers: the first row has eight blades, the second - six. 7 types of models of turbulent viscosity, which most fully describe the phenomena in the flow around the propfan and the inlet device, have been investigated: k-ωmodel; SSТ (shear stress transport) SST Transitional №1 Fully turbulence; SST Transitional №2 Specified Intermittency; SST Transitional №3 Gamma model; SST Transitional №4 Gamma theta model; SST Transitional №5 Intermittency. The results of mathematical modeling of the flow near the propfan and in the inlet device at the corresponding operating mode of the turbopropfan engine using the selected models of turbulent viscosity, the total pressure value in front of and behind the inlet device was obtained to determine the total pressure recovery coefficient in it and the value of the propfan thrust. The value of the recovery factor of the total pressure in the inlet device and the propfan thrust are compared with the flight test data of the prototype. An analysis of the comparison of the values of the total pressure recovery factor in the inlet device and the propfan thrust showed that the use of the SST Transitional №4 Gamma theta model allows obtaining the value of the total pressure recovery factor in the inlet device and the propfan thrust that is closest to the flight test results.

Parameterization and optimization design of a two-dimensional axisymmetric hypersonic inlet

Optimization method, as a promising way to improve inlet aerodynamic performance, has received increasing attention. The present research is undertaken to design a two-dimensional axisymmetric hypersonic inlet using parametric optimization. The inlet configuration is parameterized and optimized in consideration of total pressure recovery and starting performance. A Pareto front is obtained by solving the multi-objective optimization problem. Then, the flow structures of the optimized inlets are analyzed and the starting performances are evaluated. Results show that the total pressure loss mainly occurs in the internal contraction section, especially near the inlet entrance, and therefore the total pressure recovery coefficient can be greatly improved by decreasing external compression. As a result, the guidance for designing high-performance inlets is concluded. Besides, it is found that as the internal contraction ratio increases, the inlet starting ability becomes worse, which attributes to the larger separation bubble at the inlet entrance. Finally, the total pressure recovery coefficient and the starting Mach number of the optimized inlets are obtained, which can be a reference for engineering design.

Numerical analysis on the conceptual design of an air-breathing engine inlet working in mach number 3∼5.5

Based on the equal-intensity shock theory, this article designed a supersonic inlet working in Mach number 3.0∼5.5 with the background of an air-breathing engine. The inlet applied the four-shock train mixed compression configuration and inserted a sidewall compression at the beginning of the isolator. Through developing effective 3D RANS computations validated by current experiments, the performance of the designed inlet was identified. The designed inlet self-starts at freestream Mach number Ma∞ = 3.0 under which the total pressure recovery coefficient has dramatic increment, and the aerodynamic choking at the inlet throat no longer presents; the inlet keeps working at all studied flight states with zero angle of attack (AoA) and achieves shock-on-lip at the design point Ma∞ = 5.0. Both positive and negative AoAs can disrupt the equal-intensity shock allocations, which degrade the inlet performance. The inlet obtains maximum total pressure recovery coefficient at zero AoA. The maximum back pressure at Ma∞ = 3.0∼5.5 obtained by the inlet surpasses the requirements and keeps a certain margin. The inlet performance basically meets all the goals proposed by the engine.

Study on the aerodynamic performance of boundary-layer-ingesting inlet with various geometries

The mixing loss in S-shaped duct causes the contradiction between the total pressure recovery and circumferential distortion, which confuses the design process for boundary layer ingestion (BLI) inlet. This article is aiming to present a detailed understanding of the effect of two kinds of loss resources (mixing loss and frictional loss) on the performance for various BLI inlet geometries. To achieve the above objective, a design method including three control parameters is established. With this method, a systematic investigation of control parameters on the performance of BLI inlet is carried out, which leads to the contour map for the performance of BLI inlet. Based on the map, the flow physics of BLI inlet with two kinds of losses is then analyzed. Finally, an empirical design principle is provided. The results show that the appropriate design for BLI inlet has the potential to increase the total pressure recovery or reduce the circumferential distortion.

Experimental and numerical investigation of a supersonic inlet with double S-bend diffuser

Design of a supersonic inlet with double S-bend diffuser was developed. Numerical simulation and wind tunnel experiment were carried out to investigate the aerodynamic performance and variable geometric rules of the inlet. The result indicates that the variable geometry scheme adopted solves the contradiction between starting performance at low Mach number and aerodynamic performance at high Mach number. The inlet works normally and stably over a wide speed range. At design point, the total pressure recovery coefficient reaches 0.47. In addition, two different kinds of inlets with double S-bend diffuser and single S-bend diffuser were studied. Compared with the double S-bend diffuser, the total pressure recovery coefficient of the single S-bend diffuser is higher at low Mach number (Ma0 < 3) and lower at high Mach number (Ma0 > 3). With the increase of backpressure, shock train mainly moves upstream along the low-energy flow region in the diffuser. For the double S-bend diffuser, shock train will first move along the lower corner and then along the upper corner. For the single S-bend diffuser, it will only move along the upper corner. The strong secondary flow of the double S-bend is the main reason for the above phenomenon.

Development, Analysis, and Validation of a Simultaneous Inlet Total Pressure and Swirl Distortion Generator

Over the last decade, the Turbomachinery and Propulsion Research Laboratory at Virginia Tech has researched, invented, developed, computationally analyzed, experimentally tested, and improved turbofan engine inlet distortion generators. This effort began with modernizing and improving inlet total pressure distortion screens originally conceived over half a century ago; continued with the invention of inlet swirl distortion generators (StreamVanes™) made possible only through advances in modern additive manufacturing technology; and has, thus far, culminated in a novel combined device (ScreenVanes™) capable of simulating realistic flight conditions of coupled inlet total pressure and swirl distortion in a ground-test turbofan engine research platform. The present research focuses on the methodology development, computational analysis, and experimental validation of a novel simultaneous inlet total pressure and swirl distortion generator. A case study involving a single bend S-duct inlet distortion profile demonstrates the ability to generate a high-fidelity profile simulation, yet outlines a design process sufficiently generic for application to any arbitrary inlet geometry or distortion profile. A computational fluid dynamics simulation of the S-duct inlet provided the target profile extracted at the aerodynamic interface plane. Next, utilizing a method of inverse propagation, the planar distortion profile was propagated upstream to yield a flow field that could be manufactured by a distortion generator adequately isolated from turbomachinery effects. The total pressure distortion screen and swirl distortion StreamVane components were then designed and computationally analyzed. Upon successful computational reproduction of the S-duct inlet distortion profile, experimental hardware was fabricated and tested to validate the ScreenVane methodology and distortion generating device. Comparison of the S-duct manufactured distortion and the ScreenVane manufactured distortion was used as the primary criterion for profile replication success. Results from a computational analysis of both the S-duct and ScreenVane indicated excellent agreement in distortion pattern shape, extent, and intensity with full-field total pressure recovery and swirl angle profiles matching within approximately 0.80% and 2.6°, respectively. Furthermore, experimental validation of the ScreenVane indicated nearly identical full-field total pressure recovery and swirl angle profile replication of approximately 1.10% and 2.6°, respectively, when compared to the computational results. The investigation concluded that not only was the ScreenVane device capable of accurately simulating a complex inlet distortion profile, but also produced a viable device for full-scale turbofan engine ground test.

Aerodynamic and Stealth Characteristics Analysis for a New Type of S-Shaped Inlet With Inner Bulge

Searching for the optimal solution of the comprehensive performance of the aerodynamic and stealth characteristics of the aircraft inlet has become the trend of the development of the aircraft inlet due to the rapid development of radar detection technology. The aircraft inlet is a vital source of scattering due to its own cavity structure characteristic. Accordingly, reducing the electromagnetic stealth characteristic value of the stealth aircraft inlet can be served as the main way to reduce the RCS value of the whole airplane. Accordingly, a new type of stealth aircraft inlet that incorporates the design of the inner bulge is designed in this paper, which significantly reduces the value of radar-cross section (RCS) of the inlet by transforming the three-dimensional shape of the inlet. The design of this kind of inlet is based on the parametric rapid reconstruction geometry method that produces a geometry profile of the inlet through controlling a limited number of characteristic parameters. At the same time, the profile coordinates of the inner bulge is decided by means of three parameters, including the area of inner bulge, the circumferential distance and the radial offset. This paper calculates the aerodynamic characteristics of the seven S-shaped inlets, including six S-shaped inlets incorporated the inner bulges with different parameters, and an ordinary S-shaped inlet without the inner bulge as a control model. Total pressure recovery and distortion index are used as the indicators for judging the flow characteristics of these inlets to evaluate the effects on the aerodynamic characteristics of the stealth inlet after incorporating the inner bulge, which plays an important role in further evaluating the influence of the working environment of the compressor after the inlet. These S-shaped inlets with inner bulges on them resulting a decrease on total pressure recovery by 0.69% to 1.14%, at the same time, the total pressure distortion index at the exit of inlet is increased by 14.77% to 26.85%. Overall, the aerodynamic performance of the inlet with the design of the inner bulge can’t be seen a large decline. What’s more, this paper also compares the electromagnetic properties of the S-shaped inlet models with the inner bulge with the inlet without the inner bulge, using the algorithm of iterative physical optics (IPO) to evaluate the stealth characteristics of the models. This method is a kind of high-frequency approximation method for analyzing the scattering characteristics of the cavity scattering, with the high calculation accuracy as well as the high calculation efficiency. At the same time, the calculation takes up less memory, this is simply because that the mesh is coarser in calculation when using the IPO algorithm. The resulting inlet design achieving a dramatic improvement in stealth performance that the value of RCS is reduced by up to 97% after incorporating the design of inner bulge. In general, the stealth characteristics of optimized stealth aircraft inlet has greatly improved.

Blowing Control in Serpentine Inlet and its Effect on Fan-Stage Performance

Serpentine inlet is widely used in military and civil aircraft due to its good stealth performance. However, it generates a high total pressure loss and swirl distortion which significantly affects the performance and the stability of the compressor. In order to improve the quality of the flow field at the aerodynamic interface plane (AIP), a flow control is required inside the serpentine inlet. The objective of this paper was to study the effectiveness of the blowing active flow control on reducing the swirl distortion and on improving the total pressure recovery at the AIP, by reducing the low-momentum flow in the serpentine inlet. The mechanism of the blowing control and the effect of the design parameters (i.e. blowing angle, blowing position and blowing flow rate) on the aerodynamic performance at the AIP were studied. The optimal solution was applied to the full flow path of the serpentine inlet and the fan-stage. The numerical results showed that the quality of the flow field at the AIP were effectively improved by blowing high-energy airflow into the boundary layer of the serpentine inlet. The blowing position had a high influence on the blowing effect, and upper wall blowing scheme obtained greater benefits than lower wall blowing scheme and combination blowing scheme. In addition, the blowing angle should be selected to avoid the high-energy air from pipes mixing with mainstream in the serpentine inlet which will result in an additional total pressure loss. When the ratio of the blowing mass flow rate to the designed mass flow rate of the serpentine inlet was about 1.5%, the swirl distortion on the AIP reached a minimum value, which then did not show a significant difference in performance with blowing ratio increased. When the upper wall blowing scheme was adopted with a blowing angle of 6 degrees and a blowing ratio of 1.5%, the AIP aerodynamic performance achieved the highest improvement, with an increase of the total pressure recovery factor by about 1%, and a decrease of the circumferential total pressure distortion and the swirling distortion by 60% and 61%, respectively. With the optimal control scheme, the area of the low-pressure region near the upper wall was remarkably reduced, and the performance of fan-stage was improved, with an increase of the pressure ratio by about 1.5%, and the efficiency of the single-stage compressor by about 3.1%, respectively.