Increase In Total Pressure Loss Research Articles

Investigation of an Inter-Compressor S-duct using the Lattice Boltzmann Method

Abstract This article presents the study of an inter-compressor annular S-duct. Numerical simulations are performed using Large-Eddy SImulation based on a compressible hybrid thermal lattice Boltzmann method (LBM) implemented within the ProLB solver. Comparisons are made between the LES-LBM results, Reynolds Averaged Navier-Stokes (RANS) computations, and experimental measurements on a representative S-duct taken from the European project AIDA. Several cases with increasing complexity are addressed where the different rows surrounding the duct are gradually included in the computations. The effects of each row on the flow field development and loss levels are studied. The goal is to evaluate the ability of the LES-LBM to recover the aerodynamic behavior and the total pressure loss evolution within the duct. Results show that the LES-LBM retrieves the correct flow evolution inside the S-duct compared to the experiment and previous RANS results. The case where the upstream stator row or the Low-Pressure Compressor (LPC) stage is integrated shows an increase in total pressure loss, as previously observed in the literature, and a more developed flow field with complex flow features contributing to the loss generation. To further analyze the loss mechanism, an entropy-based approach is presented and highlights that most losses are generated close to the hub wall due to the migration of the upstream stator wakes.

Optimization strategies for the design of pre-cooler based on the synergistic air-breathing rocket engine

A foundation for the engineering of pre-coolers is provided to understand the flow and heat transfer characteristics of air pre-coolers. Using the pre-cooler in the combustor of the Synergistic Air-Breathing Rocket Engine (SABRE) as the research object and selecting the smallest periodic unit of the model, numerical methods were adopted to study the flow field, outlet temperature, total pressure loss, heat transfer coefficient, and heat transfer power under the different number of tube rows, pitch, and helium/air heat capacity ratios. The number of tube rows increased from 7 to 15, the tube pitch increased from 1.5 mm to 3.5 mm, and the helium/air capacity ratio increased from 1.06 to 1.64. The results show that increasing the tube pitch reduces the outlet velocity and decreases the heat exchange efficiency, but the total pressure loss is reduced. Increasing the number of tube rows results in a linear increase in total pressure loss, but the heat exchange efficiency also increases. The increasing trend slows down when the number of tube rows exceeds 11. An increase in the heat capacity ratio can reduce the total pressure loss, increase the heat exchange power, and increase the amount of helium required. It can provide a reliable technical foundation for the practical engineering design of similar pre-coolers.

Secondary flow and loss mechanisms of variable stator vanes in an annular cascade

Axial compressors with variable stator vanes require annular gaps and radial gaps from the endwalls for smooth adjustment, which induces complex secondary flows such as the penny leakage vortex and tip leakage vortex, leading to a negative impact on the aerodynamic performance. To better understand these mechanisms, numerical investigations were conducted on four different clearance configurations. The results show that the penny leakage vortex moves toward the suction side under the transverse pressure gradient and mixes with the hub corner stall vortex. This causes the corner separation to be further developed, leading to an increase in total pressure loss by 13.6%. However, the tip clearance leakage flow could reduce the transverse pressure gradient, which prevents penny leakage vortex from mixing with low-energy fluid in the corner region. Moreover, the hub corner stall vortex is also replaced by the tip leakage vortex, which effectively suppresses the range of corner separation. Under the comprehensive effects of the penny leakage vortex and the tip leakage vortex, the total pressure loss coefficient is increased only by 7.6%. Therefore, the mixing effect between the penny leakage vortex and low-energy fluid in the corner separation is the main reason for higher loss production of the cascade, and these findings provide theoretical support for the future application of flow control technology to reduce secondary flow loss.

Influence of flow control schemes of the three-strut injector on the mixing efficiency and total pressure loss for a scramjet combustor

Mixing efficiency and total pressure loss of a scramjet combustor with a three-strut injector under different flow control schemes are numerically investigated using the two-dimensional Reynolds-Averaged Navier-Stokes equations with the SST k-ω model. The flow structure is analyzed in detail to reveal the mixing enhancement mechanism of the three-strut injector configuration. The results demonstrate that the three-strut injector configuration significantly improves the mixing efficiency in the ignition region of the scramjet combustor by about 20% and only about 2% increase in total pressure loss compared to the single-strut injector configuration. The strut and nozzle size distributions of the central and auxiliary struts has negligible impact on the mixing efficiency, but they can alter the flow structure. An increase in the separation distance between the central and auxiliary struts promotes the interaction between shock waves and shear layers, resulting in a reduction of the complete mixing distance by nearly 80 mm. In addition, increasing the inward injection angle of the auxiliary struts improves the mixing efficiency of air and hydrogen in the ignition zone by nearly 15%. The rise in total pressure loss caused by the flow control schemes for the three-strut injector configuration is negligible.

Scramjet combustion with bumpy struts

In this study, the following strut configurations were proposed to improve the combustion performance of a scramjet combustor with a strut: (1) BS-1: strut with periodic rectangular bumps and (2) BS-2: strut with periodic trapezoidal and triangular bumps. These bumps were arranged on both sides of the rear part of the struts. By using a validated numerical simulation method, the performance and combustion flow fields of scramjet combustors with a normal strut (NS) and two BSs were compared and analyzed under an inlet Mach number of 2.0 and hydrogen jet at sonic speeds. The effects of the two BSs on the performance of the scramjet combustor and their mechanisms were investigated. The results demonstrated that in the two BS combustors, the flow disturbance and streamwise vortices induced by the bumps, additional oblique shock waves at the leading edge of the rear straight section, and upstream shift in the interaction position between the reflected oblique shock wave and shear layer resulted in a significant increase in total pressure loss. However, the flow disturbance and streamwise vortices not only significantly increased the shear layer thicknesses and interfacial area between the fuel and air, which contributed to a more stable and efficient combustion by increasing the size of the near-field high-turbulent-kinetic-energy region and expanding the combustion region, but also enhanced heat and mass convection between the core region and main stream on both sides, yielding a higher flame propagation speed. Therefore, the combustion was significantly promoted in the two BS combustors. BS-2 had a stronger disturbance effect on the flow and induced stronger streamwise vortices, resulting in a greater total pressure loss compared with that observed in BS-1. However, the triangular bumps of BS-2 increased the width of the narrow part of the recirculation zone, which made the fuel distribution downstream of BS-2 combustor more conducive to maintaining a larger combustion region, thus achieving more efficient combustion. Compared with the NS combustor, BS-1 and BS-2 combustors reduced the distances by 49.4% and 54.1%, respectively, when the combustion efficiency reached 95%, and by 47.9% and 53.1%, respectively, when the combustion efficiency reached 100%, with the total pressure loss at the outlet increasing by 2.16% and 3%, respectively. Therefore, the results suggest that the two BSs have the potential to significantly enhance the combustion of scramjet combustors, making them a promising avenue for further research and optimization.

Effect of different strut design on the mixing performance of H2 fueled two-strut based scramjet combustor

Strut fuel injection approach significantly influences the combustion mechanism inside the scramjet engines. To address this issue, researchers are nowadays mainly focused on a scramjet model with two struts. As the performance of the two-strut model is improved as compared to the original single-strut model as per the latest research findings, therefore the current research is focused on the impact of the different shapes of the two-strut on the combustion flow-field study of scramjet engines. Initially, experimental data from the already-published literature for a single strut is used to validate the implemented simulation techniques. Then, to investigate the combustion performance, different shapes of the two-strut are included in the already-existing DLR scramjet combustor. The SST k-ω turbulence model and steady RANS equations are used to carry out the numerical simulation. The semi-circular trailing edge half-wedge strut, with a combustion efficiency of roughly 92%, is the most effective strut configuration obtained in the current study. The findings also revealed that the semi-circular trailing edge half-wedge strut profile attains a stronger as well as a wider wake region, ensuring better combustion performance, and indicating that the two-strut model can be used as an efficient flame holder in subsequent works. Next, an analysis of how various inward fuel injection angles affect the performance of the optimum strut shape in terms of mixing efficiency and total pressure loss is provided. According to the study, an increase in the inward fuel injection angle significantly improves the mixing performance of the two-strut scramjet combustor. Additionally, the mixing and total pressure loss profiles reveal that the geometry with a 60° strut injection angle achieves early complete mixing within 0.145 m at the cost of only a slight increase in total pressure loss, and this outcome may lead to the development of a short-length supersonic combustor to minimize the drag forces and heat generation. Last but not the least, the optimal inward fuel injection angle (60°) of the two-strut and its effects on mixing performance under different freestream conditions are investigated, illustrating that mixing efficiency decreases with increasing Mach number in the same manner as single strut geometry.

Numerical Study of Film-Cooling and Aerodynamic Performance of Contoured Cratered Holes

Film cooling with discrete cratered holes has been proven to be one of the effective ways to increase the cooling effectiveness for flat-plate conditions. Up to now, there has been a lack of relevant research on various discrete cratered holes under cascade conditions. In this paper, the cooling and aerodynamic performances of a reference contoured cratered hole and two optimized contoured cratered holes are studied employing the numerical analysis method. The velocity and temperature fields for the film cooling are obtained by solving the steady-state Reynolds-averaged Navier-Stokes equations with the Shear Stress Transport turbulence closure. The numerical results show that these two optimized holes significantly enhance the cooling performance with an ignorable increment of pressure loss compared to the reference hole. When the cooling hole is placed at the pressure or suction side, the local pressure gradient plays a key role in developing the downstream flow field for either cooling hole. Furthermore, the pressure loss is mainly affected by the blowing ratio and the film cooling hole location. The pressure loss in the hole and its contribution to the total pressure loss increase with the increase of the blowing ratio.

Effect of strut profile, angled injection, and blunting of leading edge on mixing of supersonic hydrogen-air streams

Mixing fuel and air in a scramjet combustor is one of the key challenges in its development due to the short residence time available. Strut based injection is considered a viable alternative and is widely investigated. In the present work, the effect of the strut profile, angled injection of fuel, and the blunting of the leading edge of the strut on the mixing of fuel and air is studied using the Finite Volume Method (FVM). The base case chosen is an experiment with Hydrogen fuel injection from a strut at Mach 2.3 in a supersonic air stream of Mach 2. The popular strut profiles of diamond and wedge shapes are considered. The mixing due to wedge strut with dual injection from the trailing face and the effect of the inclination angle of the two fuels is studied. It is observed that the wedge strut with single fuel injection results in the highest mixing efficiency. Compared to the diamond strut, the distance to reach the complete mixing of fuel and air is reduced by 100 mm for the wedge strut with single injection. The high thermal load on the leading edge of the strut requires blunting (and rounding) for heat removal, and its effect is also investigated. The blunting of the leading edge of the strut leads to a decrease in mixing efficiency and an increase in total pressure loss as well as rise in entropy compared to a strut with pointed leading edge. Blunted wedge strut with single injection requires 40 mm more length for complete mixing than wedge strut with pointed leading edge. Also, after blunting, its rounding increases the mixing efficiency slightly but reduces the total pressure loss and entropy rise noticeably.

Numerical study of unsteady shock/mixing interaction in confined space

Unsteady shock wave enhancement is an effective technical approach to achieve air/fuel mixing under high speed and compressible conditions. However, unsteady shock/mixing interactions are always accompanied by complex wave structures, so the height of the combustion chamber will inevitably have a significant impact on shock waves and engine performance. In this regard, shock/mixing interactions in confined space are studied in this article by direct numerical simulation to investigate the effect of wall constraint on shock wave enhancement. Shock wave structure, total pressure loss, thickness of mixing layer, and turbulence intensity of flow field with different confinement degrees are calculated and compared. The results show that the unsteady shock wave can dramatically enhance the thickness of the mixing layer and the turbulence intensity, thus improving the mixing efficiency. In addition, as the wall constraint increases, the intensity of the shock wave and the total pressure loss increase; the flow field is more prone to Mach reflection. The average total pressure loss caused by steady shock wave and unsteady shock wave is essentially the same for an identical wall height, but the total pressure distortion caused by the unsteady shock wave is more substantial. Furthermore, the enhancements of mixing efficiency by the shock wave are also affected significantly by the degree of wall constraint; therefore, a suitable wall height needs to be determined in order to achieve the highest mixing efficiency.

Numerical investigation on geometric sensitivity and flame stabilisation mechanism in H2 fueled two-strut based scramjet combustor

The layout of the strut fuel injector has a significant impact on the flame stabilization process of the strut-based scramjet combustor. That’s why a two-strut-based scramjet model has been developed in recent researches. As several recent studies have shown that the two-strut model performs better than a single-strut model in terms of performance parameters, the current research thus focuses on the effect of geometric sensitivity to the combustion flow-field as well as flame stabilization mechanism of the two-strut scramjet engine. To begin with, the adopted computational approaches are validated using experimental data from the existing literature for a single strut. Following that, multiple ramp positions are imported onto the existing two-strut geometry and then divergence angles on both the walls of the two-strut combustor are also varied to explore the combustion performance. The present investigation reveals that geometric variation has a significant influence on the performance study of a supersonic combustor. Increasing the ramp position from the combustor entrance leads to a slight improvement in combustion efficiency but at the cost of a corresponding increase in total pressure loss. Moreover, increasing the divergence angle on both walls weakens the combustion performance to some extent. Finally, a flame stabilization analysis is carried out in a two-strut model, which identifies two stable modes of combustion in the same way as single strut geometry does, suggesting that two-strut geometry can be adopted as an effective flame holder in future works.

Effects of Turning Angle and Turning Internal Radius on Channel Impingement Cooling for a Novel Internal Cooling Structure

Abstract Leading-edge multi-channel double-wall design, a novel internal cooling structure, has been presented recently to enable higher overall cooling effectiveness with less penalty of coolant and pressure loss. Our previous work has proved the advantages of the design under operating condition relative to conventional internal cooling methods, including impingement cooling and swirl cooling. Channel impingement cooling structure, which is utilized at the turning region of the leading edge, is the critical factor to realize the high cooling performance of the design. Hence, the turning angle and turning internal radius of the cooling channel are two key parameters for the novel design, and this paper focuses on the effects of these two parameters on the flow and heat transfer characteristics of the design. Nine simplified single-channel models with different turning angles (45 deg, 60 deg, and 75 deg) and radii (0.6 mm, 0.9 mm, and 1.2 mm) were adopted to conduct the study, and the jet Reynolds number ranges from 10,000 to 40,000. The results show that the turning angle and turning internal radius affect the jet form significantly for the same mechanism. Small turning angle means large impingement, which leads to stream-wise counter-rotational vortices and high turbulence intensity, but increasing turning internal radius transfers the jet form from impingement jet to laminar layer attaching the target surface with low heat transfer. The turning internal radius has stronger effect than turning angle. With higher jet Reynolds number, both the heat transfer and total pressure loss increase dramatically, and the effects of geometrical parameters are clearer.

Investigation on Performance of Compressor Cascade with Tubercle Leading Edge Blade

Abstract In order to improve compressor performance using a new design method, which originates from the fins on a humpback whale, experimental tests and numerical simulations were undertaken to investigate the influence of the tubercle leading edge on the aerodynamic performance of a linear compressor cascade with a NACA 65–010 airfoil. The results demonstrate that the tubercle leading edge can improve the aerodynamic performance of the cascade in the post-stall region by reducing total pressure loss, with a slight increase in total pressure loss in the pre-stall region. The tubercles on the leading edge of the blades cause the flow to migrate from the peak to the valley on the blade surface around the tubercle leading edge by the butterfly flow. The tubercle leading edge generates the vortices similar to those created by vortex generators, splitting the large-scale separation region into multiple smaller regions.

Sweeping Jet Film Cooling on a Turbine Vane

The cooling performance of sweeping jet film cooling was studied on a turbine vane suction surface in a low-speed linear cascade wind tunnel. The sweeping jet holes consist of fluidic oscillators with an aspect ratio (AR) of unity and a hole spacing of Pd/D = 6. Infrared (IR) thermography was used to estimate the adiabatic film effectiveness at several blowing ratios and two different freestream turbulence levels (Tu = 0.3% and 6.1%). Convective heat transfer coefficient was measured by a transient IR technique, and the net heat flux benefit was calculated. The total pressure loss due to sweeping jet film cooling was characterized by traversing a total pressure probe at the exit plane of the cascade. Tests were performed with a baseline shaped hole (SH) (777-shaped hole) for comparison. The sweeping jet hole showed higher adiabatic film effectiveness than the 777-shaped hole in the near hole region. Although the unsteady sweeping action of the jet augments heat transfer, the net positive cooling benefit is higher for sweeping jet holes compared to 777 hole at particular flow conditions. The total pressure loss measurement showed a 12% increase in total pressure loss at a blowing ratio of M = 1.5 for sweeping jet hole, while 777-shaped hole showed a 8% total pressure loss increase at the corresponding blowing ratio.

Numerical investigation of droplets-gas mixing performance in depth adjustable underwater launcher cooling chamber

PurposeThe purpose of this study is to investigate structure parameters that influence the mixing process of droplets-gas in underwater depth-adjustable launcher cooling chamber and help engineers who design the launcher to distinguish the most important factor that impacts mixing performance in the cooling chamber.Design/methodology/approachEuler–Lagrangian droplet tracking method was used to simulate droplets-gas mixing process in the cooling chamber. The SST k-w model was adopted to simulate turbulence. Droplet breakup was described by KHRT hybrid model using modified contains which are more fit to the supersonic main flow condition.FindingsThe results show the counter-rotating vortex pairs which caused by injected liquid accelerate the mixing process. High-pressure supersonic freestream makes the liquid jet break into more small droplets due to the high momentum of the main stream. Axial injection angle has the greatest influence on Sauter mean diameter (SMD). Penetration height, SMD and total pressure loss slightly change in different tangential injection conditions. However, mixedness decreases with reduction of tangential injection angle due to a more limited space for spray developing. Enlarging orifice diameter raises penetration and mixedness greatly, while SMD and total pressure loss increase slightly.Originality/valueThe findings of this study confirm the key structure parameter to improve mixing performance in the cooling chamber. Engineers who design the underwater depth-adjustable launcher can refer the findings in this study to make control of launching power more accurate.

Modeling Particle Deposition Effects in Aircraft Engine Compressors

Airborne particles ingested in aircraft engines deposit on compressor blading and end walls. Aerodynamic surfaces degrade on a microscopic and macroscopic scale. Blade row, compressor, and engine performance deteriorate. Optimization of maintenance scheduling to mitigate these effects requires modeling of the deterioration process. This work provides a deterioration model on blade row level and the experimental validation of this model in a newly designed deposition test rig. When reviewing previously published work, a clear focus on deposition effects in industrial gas turbines becomes evident. The present work focuses on quantifying magnitudes and timescales of deposition effects in aircraft engines and the adaptation of the generalized Kern and Seaton deposition model for application in axial compressor blade rows. The test rig's cascade was designed to be representative of aircraft engine compressor blading. The cascade was exposed to an accelerated deposition process. Reproducible deposition patterns were identified. Results showed an asymptotic progression of blade row performance deterioration. A significant increase in total pressure loss and decrease in static pressure rise were measured. Application of the validated model using existing particle concentration and flight cycle data showed that more than 95% of the performance deterioration due to deposition occurs within the first 1000 flight cycles.

Experimental investigation of the role of struts in high speed mixing

AbstractThis paper presents the experimental results of the role of struts in supersonic mixing. Experiments were carried out with novel strut configurations to show their capabilities on mixing with reasonable total pressure losses. The performances were compared with the Baseline Strut configurations (BSPI and BSNI). The analysis presented includes the mixing quantifications using Mie scattering signature, flow field visualisation, measurement of wall static pressure and the total pressure loss calculations. The results clearly demonstrated that the proposed strut configurations achieved increased mixing (7-8%) compared to BSPI with increase in total pressure loss (2%). On the other hand, when compared with BSNI, the mixing performance was found to be decreased by 6% with reduced total pressure loss (12%).

Effects of condensing moist air on shock induced oscillation around an airfoil in transonic internal flows

Shock induced oscillation around an airfoil in transonic internal flow fields are often observed due to complex shock wave boundary layer interaction. However, in actual flow where finite amount of water vapor is present in the air, the rapid expansion of the flow may give rise to non-equilibrium condensation. Effects of condensing moist air on unsteady shock induced oscillation were numerically studied for total back pressure to reservoir pressure ratios of 0.73–0.65. Results showed that in case with condensing moist air flows, the root mean square of shock induced pressure oscillation and the oscillation frequency were reduced significantly compared with those without the non-equilibrium condensation. However, there was an increase of total pressure loss for condensing moist air flows.

Aerothermodynamics of a High-Pressure Turbine Blade With Very High Loading and Vortex Generators

Abstract AITEB-2 is a project where aerothermal challenges of modern high pressure turbine designs are analyzed. One of the scopes of the project is to allow for new gas turbine designs with less parts and lighter jet engines by increasing the blade pitch and therefore the aerodynamic blade loading. For transonic profiles, this leads to very high velocities on the suction side and shock induced separation is likely to occur. The total pressure loss increase due to flow separation and strong shocks, as well as the underturning of the flow, limits the increase of the blade pitch. In this paper, experiments using a linear turbine blade cascade with high aerodynamic loading are presented. The blade pitch is increased such that at design conditions, a strong separation occurs on the suction side. The experiments were run at high subsonic exit Mach numbers and at Reynolds numbers of 390,000 and 800,000. In order to reduce the flow separation and the aerodynamic losses, air jet vortex generators are used, which create streamwise vortices prior to the separation start. Since in high pressure turbine blades film cooling is widely used, also the influence of film cooling both with and without using vortex generators is analyzed. Film cooling is provided on the suction side by two rows of cylindrical holes. This paper provides an analysis of the influence of different main flow conditions, film cooling, and vortex generators on total pressure loss, heat transfer and film cooling effectiveness. The experiments show that the vortex generators, as well as the film cooling reduce flow separation and total pressure losses. The effects are also seen in the local heat transfer, especially with enhanced heat transport in the region with flow separation. The cases presented in this paper deal with complex flow phenomena, which are challenging to be predicted with modern numerical tools correctly. Therefore, the experimental data serve as a comprehensive database for validation of simulation tools in the AITEB-2 project.

In-Stream Hypermixer Fueling Pylons in Supersonic Flow

This paper presents results from both computational fluid dynamic and wind-tunnel experiments of in-stream fueling pylons injecting air, ethylene, and methane gas into Mach number 2.0 cold airflow. Three fuel-injection pylons studied include a basic pylon, a ramp pylon, and an alternating-wedge pylon. The latter two pylons introduce streamwise vorticity into the flow to increase mixing action. The computational fluid dynamic solution was accomplished using the commercial code FLUENT®. Three wind-tunnel experimental techniques were used: aerothermal probing, Raman spectroscopy, and nitric-oxide planar laser-induced fluorescence. Four measures reported include streamwise vorticity, total-pressure-loss, mixing efficiency, and flammable plume extent. The ramp and alternating-wedge pylons show decisive increases in mixing capability compared with the basic pylon for a finite distance downstream of the injector. The alternating-wedge pylon exhibits a measurable increase in total pressure loss compared with the basic pylon, and the ramp pylon exhibits a negligible increase in total pressure loss compared with the basic pylon. For comparison, the downstream mixing effectiveness of the three pylons is compared with the downstream mixing effectiveness of a transverse circular wall injector studied in past research. In addition, a qualitative comparison between the computational fluid dynamic and wind-tunnel experimental results is made.

Internal flow mechanism and experimental research of low pressure axial fan with forward-skewed blades

This article presents the flow mechanism analysis and experimental study of a forward-skewed impeller and a radial impeller in low pressure axial fan. The forward-skewed blade was obtained by the optimization design of the radial blade and CFD technique. Measurement of the two blades was carried out in aerodynamic and aeroacoustic performance. Compared to the radial blade, the forward-skewed blade has demonstrated the improvements in efficiency, total pressure ratio, Stable Operating Range (SOR) and less aerodynamic noise. Detailed flow measurement and computation were performed for outlet flow field for investigating the responsible flow mechanisms. The results show the forward-skewed blade can cause a spanwise redistribution of flow toward the blade mid-span and reduce tip loading. This results in reduced significantly total pressure loss near hub and shroud endwall region, despite the slight increase of total pressure loss at mid-span.