Pressure Recovery Coefficient Research Articles

Flow field reconstruction of compressor blade cascade based on deep learning methods

In the design process of compressors, calculating the S1 flow surface involves solving the Navier-Stokes equations, which results in slow convergence and makes determining cascade characteristics time-consuming. However, deep learning offers significant advantages in flow field reconstruction by not only automatically extracting complex flow features and reducing prediction time but also providing high accuracy in reconstruction. This paper implements the rapid reconstruction of the compressor S1 flow surface cascade flow field using two deep learning models: U-Net and 1D-CNN. Using a double-circular arc airfoil as an example, we selected four key design parameters that define the geometry and position of the airfoil, ultimately designing 5,292 sets of cascade geometries. By performing batch meshing and CFD simulations, we built a cascade flow field dataset. The U-Net neural network uses design parameters as input and outputs the aerodynamic distribution of the cascade flow field. After training, it can directly predict the flow field based on the design parameters. Since the U-Net model cannot directly obtain the aerodynamic parameter distribution and flow field aerodynamic coefficients on the airfoil surface, a 1D-CNN model is used as a complementary approach. The 1D-CNN model takes the design parameters as input and outputs the aerodynamic parameter distribution on the airfoil surface and the flow field aerodynamic coefficients. The prediction results show that the U-Net model achieves an average relative error of less than 1% in cascade flow field reconstruction, while the 1D-CNN model achieves an average relative error of less than 1% in predicting the pressure recovery coefficient and less than 2% in predicting the total pressure loss coefficient. This study presents a method for the rapid reconstruction of compressor blade cascade flow fields, which helps improve design efficiency and shorten the design cycle.

Analysis of the effects of combined suction on hypersonic inlet starting performance

Abstract Taking a two-dimensional inlet (Case 1) as the original inlet, the combined suction schemes are designed and their influences on the self-starting ability of the inlet are studied. The flow fields at the design point and the self-starting progress of the inlets are calculated by the numerical simulation method. The results show that the three-slot combination scheme (Case 2) leads to the reduction of the mass flow rate of the throat at the design point by 1.15% and the increase of the total pressure recovery coefficient by 6.73%. The self-starting Mach number of the inlet decreases from 4.6 to 3.2, which indicates that the self-starting ability is improved significantly by the three-slot combination scheme. In order to illustrate the role of each slot in improving the self-starting ability, three two-slot combination schemes, which are called Case 3, Case 4, and Case 5, are set up by closing Slot 1, Slot 2, and Slot 3 in turn and their self-starting processes are simulated. The results indicate that the self-starting Mach numbers of Case 3, Case 4, and Case 5 are 3.6, 4.1, and 4.5 respectively, which are 0.4, 0.9, and 1.3 higher than that of Case 2 respectively. Slot 3 has the greatest influence on the self-starting ability. When the incoming Mach number rises to 4.2, a closed separation zone is induced between Slot 2 and the throat by the reflected shock wave on the cowl side of Case 5. Due to the lack of lip overflow, which is an important flow regulation mechanism, the separation zone has a strong self-sustaining ability, and the inlet keeps unstart until the incoming Mach number rises to 4.5. In addition, in the absence of Slot 1 or Slot 2, the large-scale separation zone will not disappear until the incoming flow pressure increases to a certain value.

A reliable combustion strategy for a throttling-assisted supersonic combustor with flight Mach 7

As the flight Mach number increases, scramjet engines struggle to achieve stabilized combustion. In the present investigation, a reliable combustion strategy is developed for a kerosene-fueled supersonic combustor. Assisted by air throttling, this enables the combustor to operate efficiently at a flight Mach number of 7. Various experimental measurement techniques are used to capture the combustion process and flow characteristics. A three-dimensional numerical method based on the Reynolds-averaged Navier–Stokes equations is employed to analyse the effects of air throttling on the non-reacting and reacting flow fields. The whole combustion test is effectively divided into five processes covering the establishment of non-reacting flow, ignition, flame stabilization, and flameout. Analysis of the experimental and numerical flow fields indicates that, as the origin of the initial flame in the combustor, the aerodynamic compression induced by air throttling promotes the kerosene/air sub-mixing region. The main mixing process is enhanced by the arrangement of ramps and cavity, which contribute to effective multi-channel injection. Two combustion modes are observed, namely combined cavity shear-layer/recirculation stabilized combustion and jet-wake stabilized combustion. In the reacting flow field, the additional injection of throttling gas improves the thrust augmentation at the cost of reduced combustion intensity. The outlet combustion efficiency and total pressure recovery coefficient are found to decrease by 8.49% and 28.79%, respectively.

Design and aerodynamic performance study of subsonic Y-shaped inlet

Abstract A Y-shaped inlet with a quadrate import was designed by the super-ellipse method in this paper, and according to the variation pattern of the cross-section of the inlet, the variation pattern of the short and long axis of the super-ellipse was deduced. For many different flight situations, the numerical simulations of the Y-shaped inlet were made. The result shows (1) If 0.2≤Ma≤0.8, the total pressure recovery coefficient exceeds 0.96, and the total pressure distortion coefficient is close to 0.1, meeting the criterion of the design; (2) When Ma=0.8, for many different attack angle α=-4°∼22°, the total pressure recovery coefficient is close to 0.97, reaching the level of practical application, and it has little impact on total pressure recovery coefficient and the distortion coefficient. (3) These results demonstrate high total pressure recovery and low distortion coefficients across various flight conditions, indicating its potential for practical applications and future inlet designs.

Experimental study on the suppression of inlet blockage in rotating detonation combustor by porous-wall

The rotating detonation combustor (RDC) is renowned for its ability to provide substantial pressure gains. Nonetheless, during the stable operation of the RDC, the high-pressure rotating detonation wave (RDW) at the combustor inlet induces an increase in upstream chamber pressure, ultimately compromising engine stability and performance. To fully harness the performance advantages of the turbine-based continuous rotating detonation engine (TBCRDE) while maintaining engine stability, a porous-wall RDC has been developed to alleviate intake blockage and mitigate upstream chamber pressure rise. The operating modes, pressure rise characteristics, and performance parameters of both the porous-wall RDC and the reference configuration were systematically evaluated across varying air flow rates and nozzle designs. This analysis concentrated on the operational characteristics of the porous-wall RDC and its mechanisms for suppressing upstream chamber pressure rise. The findings reveal that the porous-wall RDC significantly extends the stable operating range and effectively reduces upstream chamber pressure rise by minimizing intake blockage. Specifically, the stable operating range is enhanced by 50 % at an outlet area ratio of 0.33, with stable rotating detonation combustion achieved at an outlet area ratio of 0.25. At an air flow rate of 1 kg/s and an outlet area ratio of 0.33, the chamber pressure rise is optimally suppressed, demonstrating a maximum reduction of approximately 16.4 %. The total pressure recovery coefficient of the combustor was analyzed, taking into account both intake loss and combustor pressure gain capabilities, and the propulsion performance of the two configurations was compared. The porous-wall RDC effectively reduces intake loss while slightly diminishing combustor pressure gain capability, resulting in a marginal increase in the total pressure recovery coefficient. Although this leads to a slight reduction in propulsion performance during chamber pressure rise suppression, the overall engine matching environment benefits from enhanced matching stability. Consequently, other engine components experience a reduced performance decline. Therefore, the implementation of a porous-wall structure is anticipated to improve the overall propulsion performance of the engine.

Transition of the flow type in the supersonic cavity controlled by the wall temperature

The wall temperature of the high-speed aircraft increases quickly under hypersonic/supersonic incoming flow, which will cause a significant change in the flow structure. To study the transition of the flow type in the supersonic cavity controlled by the wall temperature, numerical simulations are conducted. The cavity length-to-depth ratio (L/D) is varied from 10 to 15, and the wall temperature ranges from 300 K to 1300 K. The results indicate that the type of cavity flow with an L/D ratio of 13 transforms from a closed cavity flow to a transitional cavity flow, when the temperature reaches approximately 775 K. And the transitional temperature rises with the elevated total temperature of the incoming flow. Furthermore, the mechanism of the cavity flow change with wall temperature could be the competition between the recirculation zone and the shear layer in the cavity. The rising pressure with higher wall temperature in the recirculation zone weakens the downward development of the cavity shear layer, preventing it from hitting the cavity floor. As a result, the mass exchange of cavity lip surface, pressure distribution, total pressure recovery coefficient, and heat transfer distribution in the supersonic cavity change dramatically. The critical wall temperature also affected by the sidewall effects and the inflow Mach number.

Performance of an additive-manufactured precooler under high-temperature and negative pressure environment

This paper introduces an original plate-fin precooler designed for potential applications in extreme low-pressure environments. Utilizing additive manufacturing technology, the heat transfer plate thickness is constrained to 2.0 mm, with channel widths inside the plate achieving 0.8 mm. Through experimental methods, this study aims to assess the precooler’s performance and identify critical influencing factors. Experimental results indicate that the hot fluid inlet temperature to the precooler exceeds 1100 K, with static pressures dropping below 30 kPa. Despite these conditions, the precooler demonstrates an impressive pressure recovery coefficient exceeding 97 % and achieves a maximum temperature drop of 731.4 K for the hot fluid. Furthermore, it is observed that the overall performance of the precooler diminishes with increasing mass flow rates of the hot fluid, showing fluctuations of up to 25 % when assessed by the j/f1/3 factor. Additionally, while the hot fluid inlet velocity exceeds 90 m/s, laminar flow predominates during the heat transfer process. Moreover, regardless of whether the cooling fluid experiences a phase change within the precooler, its heat transfer performance show priority than that of the hot fluid. Thus, changes in the mass flow rate of the cooling fluid have minimal impact on the overall precooler performance. Finally, the first-stage heat exchanger plays a critical role in the heat transfer process, accounting for over 2/3 of the total temperature and pressure drop for the hot fluid. This research is expected to contribute to the design of high-efficiency, low-resistance precoolers, particularly those applied for operation under negative pressure conditions.

Experimental investigation of the aerodynamics of a UAV inlet with double 90° bends

Unmanned aerial vehicle (UAV) inlets commonly have a compact S-shaped inlet structure for the sake of stealth. Most of the current studies on subsonic inlets have focused on S-shaped configurations with relatively gentle transitions of the flow channel. In this study, the aerodynamic performance and swirl flow characteristics of a UAV inlet, which has double 90° bends in the duct and is integrated with an aircraft fuselage and a volute, are studied both experimentally and numerically. The influences of angle of attack, sideslip angle, AIP (Aerodynamic Interface Plane) Mach number, and freestream airspeed are analyzed. A measure of adding two deflectors in front of the first bend and a baffle at the bottom of the volute is proposed to improve the total pressure distortion and swirl distortion characteristics of the inlet. Finally, a practice of shape optimization is performed to further improve the aerodynamic performance and swirl flow characteristics on the basis of the baseline configuration. The results indicate that the maximum and minimum total pressure recovery coefficients of the baseline inlet configuration are 0.982 and 0.969, respectively, as well as the maximum total pressure distortion index of –0.0814 and the maximum swirl distortion index of 32.1 % under normal operating conditions. Through a total of 8 iterations of optimization, the total pressure recovery coefficient is eventually improved by 0.21 % of that of the baseline configuration, as well as the total pressure distortion index, swirl distortion index and maximum swirl angle reduced by 43.6 %, 4.6 % and 11.1 %, respectively.

Study on the spray characteristics of transverse jets with elliptical nozzles in supersonic crossflows using the volume of fluid–discrete phase model

The atomization characteristics of liquid jets injected transversely into a supersonic crossflow significantly affect the performance of scramjet engines. Existing research on this topic has mainly focused on circular nozzles, while the influence of nozzle geometry, particularly elliptical nozzles, has received relatively limited attention. Therefore, this study employs a numerical simulation method coupling the volume of fluid and discrete particle model to investigate the breakup and atomization characteristics of transverse liquid jets from elliptical nozzles with different aspect ratios under supersonic crossflow conditions, as well as the total pressure loss. The simulation model is validated against experimental data obtained from a pulse wind tunnel, including measurements of the liquid jet penetration depth and the Sauter mean diameter (SMD). The results indicate that for elliptical nozzles with an aspect ratio (AR) greater than 1, columnar breakup occurs earlier, and the columnar breakup length is shorter compared to circular nozzles. As the AR increases, the jet penetration depth decreases, while the spray expansion angle increases. Furthermore, the SMD of the atomized spray field from the circular nozzle is larger than that from the elliptical nozzles, and the SMD of the spray field is smallest for an elliptical nozzle with AR of 4. Finally, the elliptical nozzles exhibit a higher total pressure recovery coefficient, indicating reduced total pressure loss in the combustion chamber. This reduction in pressure loss is expected to improve the thrust performance of the scramjet engine.

Data-driven surrogate modeling and optimization of supercritical jet into supersonic crossflow

For the design and optimization of advanced aero-engines, the prohibitively computational resources required for numerical simulations pose a significant challenge, due to the extensive exploration of design parameters across a vast design space. Surrogate modeling techniques offer a viable alternative for efficiently emulating numerical results within a notably compressed timeframe. This study introduces parametric Reduced-Order Models (ROMs) based on Convolutional AutoEncoders (CAE), Fully Connected AutoEncoders (FCAE), and Proper Orthogonal Decomposition (POD) to fast emulate spatial distributions of physical variables for a supercritical jet into a supersonic crossflow under different operating conditions. To further accelerate the decision-making process, an optimization model is developed to enhance fuel-oxidizer mixing efficiency while minimizing total pressure loss. Results indicate that CAE-based ROMs exhibit superior prediction accuracy while FCAE-based ROMs show inferior predictive accuracy but minimal uncertainty. The latter may be ascribed to the markedly greater number of hyperparameters. POD-based ROMs underperform in regions of strong nonlinear flow dynamics, coupled with higher overall prediction uncertainties. Both AE- and POD-based ROMs achieve online predictions approximately 9 orders of magnitude faster than conventional simulations. The established optimization model enables the attainment of Pareto-optimal frontiers for spatial mixing deficiencies and total pressure recovery coefficient.

Influence of Double-Ducted Serpentine Nozzle Configurations on the Interaction Characteristics between the External and Nozzle Flow of Aircraft

To clarify the influence of the serpentine nozzle configurations on the flow characteristics and aerodynamic performance of aircraft, the flow features and aerodynamic performances of the double-ducted serpentine nozzles with different aspect ratios (AR), length–diameter ratios (LDR) and shielding ratios (SR) are numerically investigated. The results show that the asymmetric nozzle flow occurs due to the curved profile of serpentine nozzles, and a local accelerating effect exists at the S-bend, causing the increase in wall shear stress. The unilateral unsymmetrical expansion of the tail jet in the upward direction interacts with the separated external flow of the afterbody, forming an obvious cross-shock wave and shear layer structure. The surface pressure of the afterbody increases along the external flow direction, and decreases sharply in the separation point of the boundary layer. With the increase in AR and LDR, the local accelerating effect of the nozzle flow weakens, while with the increase in SR, the accelerating effect increases. The total pressure recovery coefficient, flow coefficient and axial thrust coefficient all decrease with the increase in AR, LDR, and SR. The thrust vector angle decreases with the increase in AR but is less affected by LDR and SR.

Effect of Fluid-Solid-Thermal Coupling on the Performance of S-Duct in Hypersonic Inlet

This article studies the effects of multi-field coupling on the S-duct by a self-developed fluid-solid-thermal bi-directional coupling code. The flow field inside the S-duct is complex with various vortices as the airflow turns. The results indicate that the total pressure recovery coefficient and the total pressure distortion index (DI) change rapidly at the beginning and reach a steady state at about 100s. The σ decreases by 12%, and the DI increases by 2% compared to the starting point. Structural deformation, size of separation zone, and wall temperature changes over time, and stabilizes around 100 s. The deformation mainly occurs at the outlet, with a maximum displacement of 3.65 mm for up-warping and 4.60 mm for expansion. The values of the total pressure recovery coefficient with S-ducts of bias distance with 20, 40, and 60 mm drop by 2.5%, 8.6%, and 12.1%, and the values of the total pressure distortion index increases with S-ducts of bias distance with 20, 40, and 60 mm increase by 10.8%, 15.1%, and 1.7%. The S-duct with a larger aspect ratio has larger separation zones and a smaller total pressure recovery coefficient.

Integrated Waverider Forebody/Inlet Fusion Method Based on Discrete Point Cloud Reconstruction

The integrated design of waverider forebodies and inlets is considered a critical challenge in high Mach number vehicle development. To facilitate the rapid construction of integrated geometrical models for waverider forebodies and inlets during the conceptual design phase, a method based on discrete point cloud reconstruction has been proposed. In this method, the geometries of the waverider body and inlet are used as inputs and decomposed into the point cloud under discrete rules. This point cloud is refitted to generate new section lines, which are then lofted into an integrated shape under the constraints of guide curves. By modifying the coordinates of the point cloud positions, the geometric configuration of the integrated shape can be rapidly adjusted, providing initial support for subsequent aerodynamic optimization and thermal protection. Using this method, an integrated approach was applied to a waverider forebody and inward-turning inlet in a tandem configuration. This achieved body-inlet matching and integration, resulting in a 15.6% improvement in the inlet’s total pressure recovery coefficient. The integration time was reduced to just 3.18% of the time required for traditional manual adjustments. Additionally, optimization based on the discrete point cloud enhanced the lift-to-drag ratio by 7.83%, demonstrating the feasibility of the proposed method.

Influence of geometric parameters on aerodynamic characteristics of serpentine convergent-divergent nozzle

The serpentine convergent-divergent nozzle represents an optimal configuration for next-generation fighter aircraft characterized by low detectability and high thrust-to-weight ratio. In contrast to the serpentine convergent nozzle, such configuration offers increased design flexibility with additional parameters, leading to heightened interactions among these parameters. As such, it is crucial to reveal the influence of design parameters on the aerodynamic performance of the serpentine convergent-divergent nozzle and the multifactor interaction, as well as its mechanism. Therefore, the influence, interaction and sensitivity of parameters on the aerodynamic performance of the nozzle were numerically investigated using the orthogonal test method. Additionally, the influence mechanism of the convergence angle, throat aspect ratio, and axial length to inlet diameter on the flow characteristics of the nozzle was investigated in detail. The results show that the convergence angle is identified as the main factor affecting the aerodynamic parameters of the nozzle. As the convergence angle increases, the thrust coefficient, total pressure recovery coefficient and discharge coefficient gradually decrease. The interaction between throat aspect ratio and other parameters is obvious. Different design parameters affect the local loss and the friction loss by affecting the curvature and wetted perimeter area, resulting in different aerodynamic characteristics of serpentine convergent-divergent nozzle.

Study on the Influence of Working Characteristics of Centripetal Pump Based on VOF/Mixture Model

Aviation fuel contamination can seriously affect aircraft flight safety, and the centripetal pump is the core component of aviation fuel purification equipment. The performance of centripetal pumps is highly demanded for purification equipment. The operating parameters of centripetal pumps significantly affect the internal flow characteristics, which affects the performance of centripetal pumps. However, the flow characteristics of a centripetal pump influenced by the operating parameters have not yet been elaborated upon. In this research, a three-dimensional numerical simulation of the air-fuel two-phase flow field inside a centripetal pump was carried out using the VOF/Mixture model to investigate the effects of three relatively independent physical quantities, namely, fuel flow, outlet fuel discharge pressure, and rotational speed, on the operating characteristics of the centripetal pump. The flow law inside the flow channel of a centripetal pump was analyzed based on a rotating fluid pressure model, the free liquid surface radius of air-fuel, and the total pressure recovery coefficient. It was found that centripetal pumps have a steady working state and an unsteady working state. In a steady working state, the proportion of separated zones in the flow channel of the centripetal pump is small, the flow coefficient C of the flow channel is greater than 1, and the total pressure recovery coefficient of the centripetal pump is high. In an unsteady working state, the separation zone in the flow channel of the centripetal pump accounts for a large proportion, the flow channel flow coefficient C is less than 1, and the total pressure recovery coefficient of the centripetal pump is low. An unsteady working state can easily occur in small flow, high-speed conditions. By analyzing the working state and flow characteristics of the centripetal pump, the mechanism of the influence of the flow, outlet fuel discharge pressure, and rotational speed on the working state of the centripetal pump is revealed, which provides a basis for the stable operation of the centripetal pump.

Aerodynamic design of supersonic compressor cascade and vorticity dynamic diagnosis of flow field structure

High-load counter-rotating compressor plays a crucial role in reducing the axial length and weight of the compressor and increasing the thrust-to-weight ratio of the aero-engine. However, the boundary layer flow separation induced by shock waves in the channel of high adverse pressure gradient also brings more aerodynamic losses. This paper proposed a supersonic compressor cascade modeling method based on the unique inlet angle theory and the superimposing thickness on the suction surface method. It carried out aerodynamic optimization design of cascade with inlet Mach number of 1.85 combined with numerical optimization technology, vorticity dynamics diagnosis, and planar cascade experiment. The results show that multiple shock wave combination pressurization can be realized in the supersonic cascade channel. At the design point, the static pressure ratio is 3.285, and the total pressure recovery coefficient reaches 86.82%, and the experimental results of planar cascade also verify the correctness of the simulation method. In addition, the correlation laws between the distribution of the vorticity dynamic parameter, shock wave structure, and aerodynamic performance of cascade were analyzed by the vorticity dynamic flow field diagnosis method, which provides a beneficial reference for the subsequent compressor design.

Experimental and numerical studies on the performance of a precooled supersonic ejector

Precooling the secondary flow enhances the performance of the supersonic ejector. However, reducing the temperature of the secondary flow without experiencing significant pressure loss presents a challenge, particularly when the secondary flow is at high temperature and negative pressure. In response to this challenge, an original design for a two-strut supersonic ejector incorporating a plate-fin precooler is developed. Experimental and numerical calculation methods are employed to analyze the system’s performance under various secondary flow conditions, including ambient air, precooled gas, and non-precooled gas. Moreover, the impact of the secondary flow, including total temperature and mass flow rate on the flow structure and mixing process of the supersonic ejector is analyzed. The results show that the shock waves, reflected shock waves and shock trains exist in the flow channel of supersonic ejector simultaneously. The shock waves intensity decreases with the rise of the total temperature and mass flow rate of the secondary flow. Moreover, this leads to the static pressure behind the nozzle changing from a ω−shape to a U-shape along the flow direction. Meanwhile, the precooler has minimal impact on the pressure loss of the secondary flow while achieving a remarkable pressure recovery coefficient of over 97% and a temperature drops of above 40%. Thus, precooling the secondary flow helps decrease the static pressure in the suction chamber, which results in a minimum 29% increase in compression ratio, enhancing the pumping capacity of supersonic ejector. The study also observed that the drop of secondary flow temperature leads to a lower velocity ratio between the secondary and primary flows, which is beneficial for their mixing process.

Design and aerodynamic characteristics of variable-geometry hypersonic inlet based on shock wave elimination

To solve the problem that the control of shock wave elimination is weakened under off-design conditions, the design concept of a variable-geometry inlet scheme that combines the variable-geometry cowl (translating and diagonalizing) with regulating shock wave elimination is introduced in this paper. The variable-geometry inlet is designed by the theories of oblique shock wave and isentropic wave as well as the Oswatitsch theory. Regulatory law of the variable-geometry cowl based on shock wave elimination is obtained by the geometric relationships between cowl compression angle, cowl shock wave angle, and optimal control point or range. Numerical simulations are conducted to investigate flow field characteristics, control mechanism, and working performance of the inlet. Results reveal that expansion waves have a significant impact on the cowl shock wave and boundary layer interaction, and flow separation. Furthermore, variable-geometry inlet with translating and diagonalizing cowl based on the regulation of shock wave elimination effectively controls and even completely eliminates the flow separation. In terms of inlet performance, the total pressure loss of the variable-geometry inlet decreases such that the total pressure recovery coefficients of the translating cowl and diagonalizing cowl inlets are increased by maximum values of 3.39 % and 9.97 %, respectively. However, the mass flow coefficient of translating cowl inlet decreases, whereas that of the diagonalizing cowl inlet is equivalent to that of the fixed-geometry inlet. The working range can be widened by changing the internal contract ratio of the inlet through translating or diagonalizing the cowl. The results confirm that the scheme of variable geometry inlet with diagonalizing cowl is practicable and reliable and has important guiding significance and value for inlet design.

Effect of strut schemes on combustion characteristics in innovative combined ramjet engine

The Mach 6-capable turbine based combined cycle engine is a potential first-stage propulsion for TSTO mission. In order to improve thrust-to-weight ratio, an innovative tandem combined ramjet engine-FABRE (Fan Augmented Air-Breathing Ramjet Engine) is designed utilizing a versatile multi-mode combustor working at both high-speed ramjet mode and low-speed turbocharged mode. Employing the strut as mixer and flame holder in multi-mode combustor presents theoretical advantages. Nonetheless, a thorough investigation into its practical feasibility and combustion characteristics is essential. This paper experimentally verifies the feasibility of strut approach in multi-mode combustor at various speeds and evaluates the effects of different strut schemes on flow field structure, thermodynamic characteristics, and combustion performance via simulation. The experimental results affirm that employing the strut scheme as a combustion organization method enables the multi-mode combustor to operate effectively under diverse inflow conditions at both high and low speeds. Simulation results further validate that these strut solutions consistently demonstrate superior combustion performance. Additionally, the influences of the strut geometry on heat release and high-temperature area are highly consistent with its effects on combustion efficiency. At low speed, wider struts lead to increased combustion efficiency, while at high speed, they initially rise before declining with further widening. Regardless of speed, increasing strut slant angle consistently reduces efficiency. At low speed, widening the strut lowers the total pressure recovery coefficient, peaking at 45° with varying strut slant angle. Conversely, at high speed, changes in strut width minimally impact, while increasing strut slant angle reduces the coefficient. These conclusions provide valuable insights for the optimization of combustion organization in the design of combined ramjet engine with a wide range of operating conditions.

Ice Object Exclusion Characteristics of Turboshaft Engine Inlet under Helicopter/Inlet Integration Conditions

In this study, the influence laws of different parameters on the exclusion characteristics of hailstone and ice flake, and on the aerodynamic performance of the inlet are studied by numerical method. The motion of the hailstone and ice flake is simulated using the 6-DOF method. Results show that the inhalation of hailstone in the inlet decreases total pressure distortion by about 20%, and the total pressure recovery coefficient is essentially unchanged. Icing of the upper lip decreases the total pressure distortion of the inlet by about 22%, and the total pressure recovery coefficient decreases by 0.6%. The ice flakes on the inner and outer lip, when shed and brake by collision with the center body, will cause damage to the engine duct. The shedding and breaking of ice flake at an angle of 150° to the lip can result in a large amount of ice flake debris entering the engine duct, threatening the performance and structure of the engine in the rear. The motion characteristics of hailstone and ice flake under helicopter fuselage/rotor/inlet integration conditions are revealed. It also provides a reference on the numerical methods for the numerical study of hailstone/ice flake exclusion characteristics of helicopter fuselage/rotor/inlet integration conditions.