Total Pressure Recovery Research Articles

Isolator shape transition impact on hypersonic internal waverider intake flow distortion

This paper examines the impact of scramjet isolator shape transition on hypersonic internal waverider (IWR) intake. The IWR intake is designed using the osculating axisymmetric flows and streamline-tracing methods. The new Internal Conical Flow “M” basic flowfield is utilized to provide the flow information for the design method. The intake is equipped with three isolators: one with a constant cross section and two with variable cross sections with circular and rectangular exits. The entrance shape and area of the three isolators are fixed to the intake throat shape and area. The exit area of the three isolators is maintained as the entrance one. Numerical computations of three-dimensional configurations reveal that the isolators with variable cross section shapes demonstrate a higher uniformity index than those with constant cross section shape. Thus, the isolator shape transition has decreased the flow distortion of the hypersonic IWR intake system. The three isolators exhibit varied wall pressure distribution depending on the isolator cross section shape, and the total pressure recovery ratios at the three isolators' exit planes are similar. The wall pressure distributions and key performance parameters at the intake throat section, including total pressure recovery, compression ratio, and Mach number, remained consistent across the first part of the intakes. Therefore, changing the cross section shape of the isolator while keeping the area constant could enhance the flow uniformity of compressed air without negatively impacting the intake system's performance. This allows a separate shape selection of the IWR intake throat and the scramjet combustor entrance to fulfill their special requirements.

Aerodynamic study of a bifurcated turboprop engine inlet with a propeller for flow at ground suction conditions

An advanced turboprop inlet with a bypass duct plays a crucial role in preventing foreign object damage and ensuring a high-quality airflow to the engine. However, it also introduces an increased flow complexity and design challenge due to the interaction between the propeller and the bifurcated ducts. To address these issues, a combined experimental and computational fluid dynamics (CFD) study was conducted on the aerodynamic performance and flowfield characteristics of a turboprop inlet equipped with a bypass duct considering the propeller interference. A ground suction test bench was utilized for generating the working conditions and the performance was measured by using total pressure rakes and pressure scanners. It is found that the rotational propeller on the one hand does work on and thus increases the total pressure recovery of the inlet, however, on the other hand causes a turning effect on the inlet flowfield structure along the direction of rotation and increases total pressure and swirling flow distortions in the engine duct. Besides, the engine duct and the bypass duct interact with each other. The combined influence of suction effect and the profile induction of the inlet leads to the majority of the shed vortices being drawn into the engine duct. Lastly, the presence of deflectors installed in the engine duct is found to effectively mitigate the secondary flow, thereby reducing the swirl distortion within the engine duct. This study may provide a significant reference to the design and optimization of advanced turboprop inlets with bypass ducts.

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.

Controlled Flow in a Serpentine Diffuser with a Cowl Inlet

Total-pressure losses and distortion in a serpentine diffuser with a cowl inlet are investigated experimentally and computationally over a range of flow rates (Mach numbers at aerodynamic interface plane MAIP of up to 0.6). The present investigations show that, in the presence of the cowl, the diffuser flow is dominated by two pairs of counter-rotating streamwise vortices that are shed from the swept edges of the cowl lip. The subsequent evolution and unsteady interactions of these vortices result in significant total-pressure losses that are coupled with high dynamic distortion at the aerodynamic interface plane. These losses and distortions are strongly mitigated by deliberately drawing ambient air across the cowl surface to form jets that interact with the cowl flow along its inner surface. The jets are formed through spanwise slots across the surface of the cowl by exploiting the pressure difference effected by the cowl flow. The interaction of these jets with the cowl flow alters the formation and evolution of the streamwise vortices, suppressing their interactions and thereby significantly diminishing the losses and distortion across the full operating range of the diffuser. For example, at MAIP=0.5, the total-pressure recovery increases by 8% while the peak circumferential distortion decreases by 35%.

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.

Pressure feedback system for flow separation mitigation in scramjet intakes

Shock Wave Boundary Layer Interactions (SWBLI) occur due to the convergence of a shock wave and a viscous boundary layer. The resulting separation bubble is a major setback for the performance of high-speed air intakes. This paper presents a numerical investigation into the potential of a Pressure Feedback Technique (PFT) to alleviate shock-induced flow separation within a Scramjet engine intake operating at Mach 4. The PFT is a novel self-sustaining flow control technique that combines simultaneous suction and injection. The two main output parameters used to support the hypothesis of the present study are the size of the separation bubble as well as the total pressure recovery at the isolator outlet. The most prominent observation of this study is that with the installation of the PF tubes, the total pressure recovery at the exit is noted to rise. The effectiveness of the pressure feedback technique in reducing the intensity of the separation bubbles is found to depend on the diameter of the PFT, pitch-to-diameter ratio, and PFT tube design.

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.

Improved Design Method for Hypersonic Intakes Using Pressure-Corrected Osculating Axisymmetric Flows Method

The efficient design of air intake is crucial for high-speed airbreathing propulsion systems. This paper presents a novel design method that integrates the recently introduced internal conical flow M (ICFM) basic flowfield (Musa et al., “New Parent Flowfield for Streamline-Traced Intakes,” AIAA Journal, Vol. 61, No. 7, 2023, pp. 2906–2921) with the pressure-corrected osculating axisymmetric flows and streamline tracing techniques. The new design method takes into account the effects of lateral pressure gradients by incorporating pressure corrections into the original design method using crossflow velocity information from the ICFM flowfield. Additionally, new entrance and throat profiles are introduced for designing two hypersonic internal waverider intakes with shape transition. One intake is designed using the new method, and the other using the original method. Viscous simulations have been performed at design (i.e., Mach 6.0) and off-design (i.e., Mach 4.0 and 3.5) conditions. The results indicate that the pressure-corrected intake exhibits superior performance in terms of total pressure recovery and reduced flow distortion. This reveals that the new design method can achieve a smooth shape and efficient aerodynamic transitions between the intake entrance and throat. The new entrance and throat profiles realized better performance than the typical ones. The startability analysis reveals that the Van Wie curve controls the considered intakes. Thus, combining the pressure-corrected osculating axisymmetric flows with the ICFM flowfield enhances the performance of hypersonic intakes, making this approach a promising avenue for future hypersonic vehicles.

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.

4D printed NiTi variable-geometry inlet for aero engines

ABSTRACT In modern aerospace engineering, the demand for innovative and adaptable solutions is paramount. This study focuses on enhancing aircraft engine components, specifically variable-geometry inlets, to meet evolving aviation technology requirements. We utilise custom Ni51Ti alloy powder to create 4D printed variable-geometry inlets. A comprehensive examination of thermal history and residual stress reveals differences in the behaviour of NiTi alloys at various points of the 4D printed inlet. After post-treatment, the printed inlet demonstrates stable two-way shape memory effects, undergoing adaptive deformation with temperature changes, mimicking the operational conditions of a commercial aircraft. Differential scanning calorimetry (DSC) and dynamic mechanical analysis (DMA) tests confirm that the 4D printed NiTi alloys maintain very stable phase transformation temperatures and two-way shape memory performance. Fluid dynamics simulations further reveal that the variable-geometry inlet design is potentially more efficient in certain operational scenarios with a total pressure recovery factor of 0.9919.

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.

Effect of tail flare of integral support plate on double-wall cooling performance

The tail of the integrated support plate replaces the traditional flame holder to stabilize the flame. To prevent the support plate from being ablated by high-temperature flame, a double-wall cooling structure is set at its tail, and cold air is introduced to cool it. However, the cold air reduces the stability of the flame. To improve the flame stability of the integrated support plate, the rear of the support plate is expanded. The effects of expansion angle on the aerodynamic characteristics of the afterburner and cooling performance of the double-wall are investigated by using the experimentally verified CFD method. The results show that the total pressure recovery factor of the integral afterburner decreases with the increase of the flare angle under the condition of non-afterburning; the flame stability is improved by increasing the flare angle within a certain range; the film covering effect at the upper part of the flame stabilizer gets worse as the flare angle increases, and the overall cooling effect of the double-wall gets worse.

Using Ultrasound Imaging To Identify Flow Reattachment In Shear-Thinning Suspensions

The current study explores ultrasound imaging velocimetry (UIV) as an alternative method for capturing flow fields in shear-thinning suspensions. In particular, UIV is used here together with pressure measurements to identify the reattachment point and to correlate the point's location with the pressure recovery downstream of an axisymmetric expansion. Four fluids are investigated: pure water, as well as three 1750ppm xanthum-gum-in-water solutions mixed with non-reactive mineral micro-particles at 0%, 15% and 30% volume fractions. Wall-pressure measurements were collected through taps located at 0 to 13 throat diameters downstream of the expansion with subsequent UIV measurements collected from 0.5 to 5 throat diameters downstream of the expansion. Pressure-tap measurements for the tested suspensions demonstrate 80% to 100% total pressure recovery achieved at approximately 4 to 5 throat diameters downstream from the throat. Furthermore, all four fluids exhibit total pressure recoveries within 10% of Borda-Carnot prediction. Despite similar pressure recoveries, UIV revealed different reattachment lengths for the tested fluids. Water exhibited a reattachment length of approximately four throat diameters, which is expected within the tested turbulent regime, while the xanthum-gum solution exhibited a delayed reattachment beyond the field-of-view when the effective Reynolds number is O(100). Pipe-wall velocity measurements suggest that the dense suspensions exhibit similar reattachment lengths to that of water despite expectations that the dense suspensions would behave similar to the pure xanthum-gum solution given: the similar Reynolds number between the pure xanthum-gum solution and the tested dense suspensions; and previous work done by Barnes et al. (2024) and others revealing shear-layer stabilization within dense suspensions. It is suspected that increased diffusion within the suspension mitigates the shear layer's ability to convect downstream, but improved measurements are required to state this conclusively.

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.