Total Pressure Loss Research Articles

Effect of Baffle Board on Aerodynamic and Stealth Performance of Double S-Duct Caret Intake

Intake is not only the main air supply component of an aircraft, but also one of the forward radar scattering sources. The aerodynamic and stealth performance of intake is critical to the serviceability of advanced fighter aircrafts. The effects of baffle boards with different configurations on the performance of the caret intake with a double S-duct diffuser are presented in this article. The multi-level fast multipole method (MLFMM) and the SST k-ω turbulence model were respectively used to calculate the surface current and the flow field. It was found that the average RCS value of intake can be effectively reduced by installing the baffle board with vertical orientation in the front diffuser, with the DC60 value and the loss of outlet total pressure both increased slightly. The boundary layer separation and the RCS characteristics of intake were closely related to the configuration of the corrugated baffle board. Compared with the traditional curved board, by installing the corrugated board with optimized corrugation number and shape, the stealth performance of intake can be further improved, and the loss of aerodynamic performance can be also reduced.

Corner Separation Control Using Sweeping Jet Actuator in a Compressor Cascade

Abstract In the pursuit of advancing active flow control (AFC) technology, a more promising non-steady actuator known as sweeping jet actuator (SJA) has been equipped on the endwall to manage the corner separation in a compressor cascade, while the steady jet actuators with holes were used for comparation. Experimental investigations have been conducted in a low-speed wind tunnel with an inflow Mach number of 0.23. Five-hole probe measurements and the oil flow visualizations were carried out to demonstrate the performance and physics of steady and unsteady blowing on controlling the corner separation. The transient data at the measurement plane were also obtained using the dynamic pressure sensors to get insight into the unsteady characteristics. Variation of jet momentum of both actuators and the excitation frequency of the SJA allows determination of favorable control parameters. Using the SJA with only 0.13% of the cascade mass flow, the total pressure loss which takes the additional energy input into account is reduced by 14.2%, while the steady jet achieves a reduction of 8.4%. The results highlight the superior characteristics of the SJA in controlling the corner separation.

Numerical study on cooling characteristics of turbine blade based on laminated cooling configuration with clapboards

The double-wall configuration, characterized by its notable overall cooling effectiveness, serves as a pivotal reference configuration in the design of next-generation turbine blades. In this study, a laminated cooling configuration with clapboards is developed, based on insights derived from both double-wall cooling configuration and laminated cooling configuration. Through a comparative analysis, with the double-wall cooling configuration and laminated cooling configuration as benchmarks, the newly constructed configuration is capable of enhancing overall cooling effectiveness while maintaining control over total pressure loss in the study. Specifically, in the context of external film cooling with an adiabatic condition, the laminated cooling configuration with clapboards attains the highest value at low blowing ratios, while the laminated cooling configuration excels in average adiabatic film cooling effectiveness at high blowing ratios. Importantly, the inclusion of clapboards significantly improves the internal heat transfer area, resulting in a greater internal total heat exchange in the laminated cooling configuration with clapboards. Consequently, this proposed configuration achieves the highest overall cooling effectiveness, achieving increases of 9.02%–14.08 % and 2.77 %–3.54 % relative to the double-wall cooling configuration and laminated cooling configuration, respectively. A comprehensive performance evaluation criterion is applied to assess cooling effectiveness and pressure loss. The results indicate that the laminated cooling configuration with clapboards enhances heat transfer effectiveness while effectively controlling pressure drop, with an evaluation value of over 1.0. Furthermore, the overall cooling effectiveness of the three cooling configurations is compared at two Reynolds numbers. It is found that the proposed configuration exhibits the best cooling performance, especially at the high Reynolds number.

Numerical studies on shock wave and boundary layer interaction in the high-load turbine rotor cascades

Design of transonic high-load turbine is the essential approach to improve thrust-to-weight ratio of gas turbine engines. The shock wave and boundary layer interaction (SWBLI) in high-load turbine is one of the important unsteady sources and the main source of loss in turbine stages. In order to study mechanism of SWBLI in high-load turbine, rotor blades with loading coefficients range from 1.6 to 2 were designed and Delayed Detached Eddy Simulations (DDES) were carried out. The influences of loading coefficient, exit isentropic Mach number and incidence angle on characteristics of shock wave, SWBLI as well as flow details in the blade passages were compared. Results indicated the shock wave was generated and enhanced as exit isentropic Mach numbers increased leading to the increase of 2.8% in the total pressure loss coefficient. There was not significant difference in the total pressure loss coefficient as loading coefficient increased, however the shock wave was stronger and the separation bubble was longer at the loading coefficient of 1.6. Meanwhile, the influence of strong separation near the leading edge of suction surface induced by variation of incidence angles on the characteristics of the shock wave and the loss was also obvious.

Action Mechanism of Film Extraction on Channel Impingement Cooling for Blade Leading Edge Using Large Eddy Simulation

Abstract Various internal cooling techniques combined with external film cooling are applied to cool gas turbine blades, and the coolant extraction can cause variations in internal flow structures and cooling effectiveness. Effects of film extraction under different film hole diameters and coolant mass flow ratio on heat transfer and flow characteristics of channel impingement cooling are presented using large eddy simulation (LES) in this paper. The current work was undertaken based on turbine blade dimensions and turbine operating conditions. The results indicate that film extraction coupled with the curvature-induced instability dominates flow patterns and heat transfer in the channel impingement cooling structure with a film hole, especially in the bend region corresponding to the leading edge of turbine blades. Film coolant extraction can disrupt the streamwise counter-rotating flow circulation pair. The latter is caused by the curvature-induced instability and is also the major driver of multi-longitudinal vortices. Hole edge vortex generated by coolant bleed can amplify heat transfer at the leading edge, which is more significant as increasing the film hole diameters. Conversely, cooling effectiveness downstream of the cooling channel would decrease due to a reduction in cooling air. Film bleed flow would increase total pressure loss, while flow loss in the cooling channel decreases because of the weaken flow circulations and less coolant. This work provides an in-depth insight into the cooling performance of channel impingement cooling with film extraction, contributing to designing film cooling for turbine blades with multi-channel wall jet cooling.

A study of angle parametric optimization of the composite honeycomb on turbine blade tip

To reduce the secondary flow loss of the high-pressure turbine caused by the tip leakage flow, the design of the composite honeycomb tip has been experimentally and numerically examined in this study. The arrangement angle of the composite honeycomb is defined and optimized using the radial basis function and genetic algorithm. The tip flow field of the high-pressure turbine blade with a 2.128 mm clearance is simulated by CFX 18.0, and the arrangement angle of the composite honeycomb is optimized to obtain a lower loss. Moreover, the flat tip, basic honeycomb tip, and optimized honeycomb tip are tested in a low-speed cascade test facility. The results show that the vortices and radial velocity induced by the honeycomb structure cavity increase the kinetic energy loss of the leakage flow and reduce the leakage flow rate. Compared with the basic honeycomb, the optimized honeycomb has a supplemental blocking effect on the tip leakage flow by reducing the crosswise pressure gradient in the blade tip. The optimized honeycomb also changes the outlet angle of the leakage flow in the tip clearance, reduces the leakage vortex, and the total pressure loss is further reduced by 2.5%.

Endwall Contouring for Improving Aerodynamic Performance in a High-Pressure Turbine Cascade

Abstract The endwall contouring has proven to be an effective technique in controlling the impacts of secondary flow within turbomachinery. A baseline cascade with the original axisymmetric endwall (BASE) has been redesigned with a non-axisymmetric contoured shroud endwall (EW-B1), and both configurations are investigated through experimental and numerical methods. The endwall is profiled through the B-spline surface method with the purpose of reducing the overall total pressure loss at the exit. The experimental studies involve flow field traverses at the exits of both cascades. Numerical calculations are conducted to gain a deeper understanding of the effects of endwall contouring. The numerical results exhibit good agreement with the experimental results, both in loss and flow angle. The results demonstrate a significant reduction in losses and secondary kinetic energy in EW-B1. The causes of the high-loss region and the overturning near the shroud are analyzed using computational fluid dynamics (CFD), as well as the changes in pressure fields and vortex structures within the cascade passage. The weakening of the horseshoe vortex and the radial movement of the passage vortex are confirmed to be the primary reasons for losses reduction.

Experimental and numerical study on flow mixing and combustion characteristics of a novel multibypass integrated combustor

The integrated design of multibypass augmented/ramjet combustors can reduce the weight of turbine-based combined cycle (TBCC) engines and improve the thrust-to-weight ratio, but low-resistance mixing and efficient stable combustion of multiple airflows over short distances are necessary prerequisites. In this study, a novel structure for a TBCC multibypass integrated augmented/ramjet combustor is proposed. The influence of the inlet aerodynamic parameters on the flow field, mixing efficiency, flow loss, and combustion performance of the combustor under different working modes was obtained via experimental and numerical methods. The experimental results show that the outlet mixing efficiency is greater than 86% in the double-bypass mode (DB-mode). While the triple-bypass mode (TB-mode) has a larger decrease, the total pressure loss is slightly reduced (by approximately 0.5%). The opening of the ram duct has a significant impact on the flow field, resulting in different rules for the influence of the inlet temperature on the outlet mixing efficiency: in the DB-mode, the mixing efficiency decreases with increasing inlet temperature, while the rule is completely opposite after entering the TB-mode. Because the dominant role of the two mixing zones in the flow field changes with the velocity, the inlet velocity has a significant impact on the mixing efficiency in the flow direction. The combustion simulation results show that the combustion efficiency in the DB-mode is almost always above 90% and the high-temperature zone is mainly concentrated downstream of the integrated strut. The radial temperature gradient increases and the combustion efficiency decreases in the TB-mode.

Investigation of control effects of end-wall self-adaptive jet on three-dimensional corner separation of a highly loaded compressor cascade

To overcome the limitations posed by three-dimensional corner separation, this paper proposes a novel flow control technology known as passive End-Wall (EW) self-adaptive jet. Two single EW slotted schemes (EWS1 and EWS2), alongside a combined (COM) scheme featuring double EW slots, were investigated. The results reveal that the EW slot, driven by pressure differentials between the pressure and suction sides, can generate an adaptive jet with escalating velocity as the operational load increases. This high-speed jet effectively re-excites the local low-energy fluid, thereby mitigating the corner separation. Notably, the EWS1 slot, positioned near the blade leading edge, exhibits relatively low jet velocities at negative incidence angles, causing jet separation and exacerbating the corner separation. Besides, the EWS2 slot is close to the blade trailing edge, resulting in massive low-energy fluid accumulating and separating before the slot outlet at positive incidence angles. In contrast, the COM scheme emerges as the most effective solution for comprehensive corner separation control. It can significantly reduce the total pressure loss and improve the static pressure coefficient for the ORI blade at 0°-4° incidence angles, while causing minimal negative impact on the aerodynamic performance at negative incidence angles. Therefore, the corner stall is delayed, and the available incidence angle range is broadened from −10°-−2° to −10°-4°.This holds substantial promise for advancing the aerodynamic performance, operational stability, and load capacity of future highly loaded compressors.

Experimental study on multi-angle pulsed jet actuators for controlling corner separation flow in a compressor cascade

Three-dimensional corner separation caused by the aerodynamic design of high-load compressor cascades is the main factor that limits their performance enhancement. Because of the unsteady characteristics of large-scale flow separation, injecting unsteady jets into the flow field has become a new flow control method. In this study, the essential flow characteristics of a compressor cascade were obtained through numerical simulations. Subsequently, the effect of pulsed jet angle on the aerodynamic performance of the cascade was experimentally investigated under different operating conditions using pulsed jet actuators. The results showed that at the design incidence angle of 0° and near-stall incidence angle of +4°, the starting point of the corner separation is located at 21% and 9% of the axial blade chord, respectively. The pulsation frequency downstream of the cascade trailing edge is 1454.5 Hz. Under the design conditions, the 20° pulsed jet exhibits the best control effect, reducing the total pressure loss coefficient by 13.4% compared with that of the original cascade. Moreover, pulsed jets demonstrate strong applicability under variable operating conditions. At near-stall incidence angles, when the jet angle is 30°, the loss-reduction ability of the pulsed jet is maximized, decreasing the loss by 25.2%. Pulsed jets reduce the cascade loss primarily by suppressing the secondary flow at the endwall and discretizing the vortex structures in the channel.

Plate-fin heat exchanger optimal design for industry employing both single and multi-objective chaotic opposition based Kho-Kho algorithm

Manufacturers of heat exchangers are constantly looking for new and better designs because it is essential to the preparation of any engineering gadget, like a Plate-fin heat exchanger (PFHE). To find an optimal design of a plate-fin heat exchanger, both hydraulic and economic issues are considered by proposing an optimization technique. The proposed technique minimizes both single and multi-objective functions like total heat transfer area, total pressure loss function, and the total annual cost of a PFHE. This work proposes a Chaotic Opposition-based Kho-Kho optimization technique to deal with single and multi-objective functions. Various benchmark functions have been assessed to evaluate the capabilities of the proposed techniques, including the Wilcoxon signed-rank test. Finally, the outputs such as total heat transfer area, total pressure loss function, and total annual cost are minimized by the proposed optimization technique, resulting in values of 78.50 m2, 0.051, and 834.64$, respectively.

Improving the combustion of scramjet engines with struts using grooves and bumps

Grooves and bumps are arranged on both sides of a normal strut (NS) to fabricate vortex generators, and introduce streamwise vortices. Struts with sawtooth grooves (GS-1), rectangular bumps (BS-1), and trapezoidal and triangular bumps (BS-2) have been proposed. In this study, struts with V-shaped grooves (GS-2) and trapezoidal and triangular bumps (BS-3) were configured. The flow fields in the cold and hot states of scramjet combustors with NS, GS-1, GS-2, BS-1, BS-2, and BS-3 were simulated numerically using steady RANS equations and an SST k-ω turbulence model. The results show that proper arrangement of the grooves and bumps on the strut can improve the mixing and combustion performance. Compared with the NS, for GS-1, GS-2, BS-1, BS-2, and BS-3, the distance required to achieve 100% fuel mixing efficiency is shorter by 21.2%, 9.1%, 18.2%, 12.1%, and 15.2%, respectively, and to achieve 100% combustion efficiency, the distance is shorter by 42.7%, 46.9%, 47.9%, 53.1%, and 58.3%, respectively. Therefore, improvement in combustion performance does not positively correlate with improvement in the mixing performance. Large-scale heat and mass convection in streamwise vortices play a dominant role in mixing. The larger the proportion of the fuel area affected by the streamwise vortices, the more conducive it is to achieving a higher mixing performance, whereas the wall has a negative impact. However, the large-scale heat and mass convection in the streamwise vortices and small-scale heat and mass convection in the high-turbulence zones are closely related to the supply of oxygen and fuel required for the combustion zone. Large-scale streamwise vortices combined with thick high-turbulence zones are conducive to achieving a high combustion rate. When the recirculation zone has a larger combustion area and smaller wall effects, a higher combustion performance is easily achieved. Compared with the grooves, the stronger disturbance effect of the bumps on the fluid effectively increased the thickness of the high-turbulence zone, thus obtaining a higher combustion performance. However, the greater blocking area and new oblique shockwaves generated by the bumps induced a greater total pressure loss. Therefore, combining bumps and grooves can further optimize strut configurations.

Multi-objective design optimization of a transonic axial fan stage using sparse active subspaces

In this paper, a multi-objective optimization strategy for efficient design of turbomachinery blades using sparse active subspaces is implemented for a turbofan stage design. The proposed strategy utilized sparse polynomial chaos expansion on a limited dataset to generate a function from which the differential and the covariance matrix can be obtained. Active subspace was used to compute the active variables via singular value decomposition and a hybrid polynomial correlated function expansion was used to construct a surrogate model on the active subspace. Coupled with freeform method and multi-objective genetic algorithm, an automated optimization loop was run at a single operating condition. An improvement in stage efficiency and total pressure ratio of 2.97% and 1.15% was achieved for the optimum design compared with the baseline. Additionally, total pressure loss coefficient decreased by 5.88%, exit flow angle by 34.65% and shock strength by 5.32%. The coupled effect of change in stagger angle, forward sweep, forward lean, and chord length reduced the recirculation at the hub, and blockage at the shroud due the tip leakage flow by decreasing the blade loading. The threshold value hyperparameter was found to be the most influential and must be accurately determined.

Optimization and experimental study of stationary endwall of stator labyrinth cavity in a low-speed research compressor

Controlling the leakage flow of the labyrinth seal cavity is very important for the efficiency and aerodynamic matching design of aero-engine high-pressure compressor. However, due to the complexity of the geometry and internal flow of the labyrinth seal cavity, the design optimization of the labyrinth seal cavity is rarely studied. In this paper, a multi-objective design optimization method was used to optimize the stationary endwall of the shrouded stator labyrinth cavity, and the optimal endwall shape was obtained. Based on the optimal geometry, the 3D printing and aerodynamic testing of experimental parts were carried out, and the aerodynamic performance and flow field details of the optimal geometry were analyzed by unsteady numerical simulation. The experimental results show that for the optimal geometry, it has a little effect on the total pressure ratio, but the efficiency at the design point is increased by 1.4 %. The numerical results show that the optimal cavity shape reduces low-energy fluid, thus reduces the total pressure loss of S3. At the near stall point, the optimal S3 still has a better performance, which means that the optimal shape is not limited to the design point. From the result, an engineering design criterion can be concluded. It is suggested that the labyrinth seal cavity should be designed to match the velocity components of the leakage flow with the mainstream of the stator properly to make the leakage flow go evenly in the blade passage instead of gathering to the suction side.

Influence of passive strut on the mixing and combustion performance enhancement

The scramjet engine is one of the possible means of reaching supersonic speed using airbreathing engines. As the inflow of air is at supersonic speed, proper mixing and combustion is a challenge. Researchers have used multi-strut arrangements to address the problem. However, the additional supply of fuel from multiple struts with inefficient mixing and combustion leads to losses in terms of fuel and efficiency. Alternatively, the multi-strut configuration could be used to improve the combustion and mixing efficiency passively. In the present investigation, two strut configurations, a fuel injection strut, and a passive strut, have been used. The computational study was done by solving steady Reynolds averaged Navier-Stokes (RANS) equations along with the Shear Stress Transport (SST) k-ω turbulence model, species, and ideal gas equation. Eddy dissipation model was used to solve reacting flow. The introduction of a passive strut in the wake of the fuel injection strut leads to vortices upstream and downstream of the passive strut. These vortices were observed to be responsible for enhancing the mixing and combustion. It was found that judicious placement of the passive strut in the wake of the combustion zone can enhance both the mixing as well as combustion efficiency with minimal loss in total pressure.

Effects of Perforated Plates on Shock Structure Alteration for NACA0012 Airfoils

This research investigated a passive flow control technique to mitigate the adverse effects of shock wave–boundary layer interaction on a NACA 0012 airfoil. A perforated plate with a strategically positioned cavity beneath the shock wave anchoring spot was employed. Airfoils with perforated plates of varying orifice sizes (ranging from 0.5 to 1.2 mm) were constructed using various manufacturing techniques. Experimental analysis utilized an “Eiffel”-type open wind tunnel and a Z-type Schlieren system for flow visualization, along with static pressure measurements obtained from the bottom wall. Empirical observations were compared with steady 3D density-based numerical simulations conducted in Ansys FLUENT for comprehensive analysis and validation. The implementation of the perforated plate induced a significant alteration in shock structure, transforming it from a strong normal shock wave into a large lambda-type shock. The passive control case exhibited a 0.2% improvement in total pressure loss and attributed to the perforated plate’s capability to diminish the intensity of the shock wave anchored above. Significant fluctuations in shear stress were introduced by the perforated plate, with lower stress observed in the plate area due to flow detachment from cavity blowing. Balancing shock and viscous losses proved crucial for achieving a favorable outcome with this passive flow control method.

Effects of wall cavity and fuel injection pressure on the performance of a non-reacting supersonic combustor

This study numerically investigates the flow field of a non-reacting cavity-configured scramjet (Supersonic Combustion Ramjet) combustor at various fuel injection pressures by solving the 2D Reynolds-Averaged Navier-Stokes (RANS) equations, species transport equations, and Menter SST k-ω model. The aim of this research is to reveal the effects of wall cavity insertion and fuel injection pressure (FIP) on the crucial performance parameters i.e., fuel-air mixing efficiency (MxE), total pressure recovery (TPR), and mass-averaged Mach number (MAMN). Accordingly, two trapezoidal cavities of aspect ratio 7 are introduced on the opposite walls of a rectangular combustor. The combustor entrance is configured with rearward-facing steps and it intakes finite parallel air streams through finite-width inlets. Gaseous hydrogen jets are injected 30 mm downstream from the combustor entrance and 10 mm upstream from the cavity leading edge. FIP is varied according to the fuel-to-freestream pressure ratios (FFPR) of 4.5, 9.0, 13.5, and 18.0. The results of the cavity-configured combustor are then compared with the performance of the combustor in the absence of the wall cavities. The results delineate the change in flow structures with the inclusion of wall cavities and variation in FIP. Insight physics of mixing, total pressure loss, and MAMN in different regions of the combustor are studied and the results are quantified for comparison. MxE in a cavity-configured combustor does not monotonically increase with decreasing FFPR as found in the combustor without wall cavities. The shock-shear layer interactions (SSLIs) play a dominant role in mixing inside the cavity-configured combustor. The results also demonstrate that the insertion of wall cavities can increase fuel-air MxE through the formation of cavity recirculation zones. In the cavity-configured combustor, a maximum of 45% MxE is achieved for FFPR 4.5, which is 4% higher than the value obtained from the combustor without the cavities with an expense of 3% greater total pressure loss.

Intake Characteristics and Inhomogeneity of Supersonic Passage

The type, quantity, location, geometry and size of the intake passage will directly affect the quality of the engine, and then affect the whole propulsion system and even the performance of the aircraft. In this paper, the intake characteristics and inhomogeneity of three different passage structures are studied under supersonic conditions. The numerical simulation models of the intake passages of F-16, F-18 and J-10 were established, and the steady-state calculations were carried out under the condition of Mach number 1.2. In the three intake passages, the total pressure decreases gradually along the flow direction, and the pressure loss near the wall is large, while the total pressure value far from the wall is large, and the pressure field is relatively uniform. The F-18 intake passage has relatively small inlet pressure and total inlet and outlet pressure loss. In terms of inlet uniformity, the intake passage of F-16 has the highest value, F-18 has lower value, and J-10 has the lowest level of inhomogeneity.

Extraction of geometric features and analysis of flow mechanism of high loaded compressor airfoils at low Reynolds number

When the aircraft cruising at high altitude, the aerodynamic performance of the compressor sharply decreases due to the enhanced boundary layer separation loss. Therefore, it is of great significance to design the high-performance airfoils suitable for low Reynolds numbers (Re) to improve the aerodynamic performance of compressor. In this paper, numerical simulations were carried out on a high-loaded subsonic compressor cascade V103 at low Re. Meanwhile, a multi-objective optimization of blade profile was performed to get optimal solutions, and data mining techniques were applied to extract valuable design knowledge from the optimal database. The results show that the Pareto-optimal airfoils achieve a maximum reduction of 27.32% in total pressure loss coefficient under design condition and 36.71% under near stall condition. Compared with the original airfoil, the curvature distribution law of the camber line with larger values at both ends and smaller values in the middle is advantageous for the performance of the compressor cascade under design condition. This distribution pattern causes a forward shift for the loading distribution of airfoils, accompanied by an earlier transition onset. Additionally, it reduces adverse pressure gradient at the rear part of suction surface, which suppresses the development and growth of laminar separation bubble and delays the separation of turbulent boundary layer. Under near stall condition, maintaining larger leading edge angle and uniform flow diffusion is beneficial for improving the performance of compressor cascade. This is because the larger leading edge angle reduces incidence angle, while the uniform flow diffusion inhibits a great flow separation.

Effect of pressure-side tip winglet with different heights and lengths on clearance flow in a compressor cascade

The clearance flow influence law and mechanism of pressure tip winglets with different fusion heights and lengths are investigated by numerical simulation methods based on a compressor cascade to weaken compressor clearance leakage and improve cascade aerodynamic performance. The result demonstrates that the tip winglet can improve the spanwise pressure flow and narrow the tip separation range, thereby suppressing the intensity of the leakage vortex. The enhanced effect on the leakage vortex is more significant when the tip winglet's fusion length range can include the leakage vortex's initial position. In addition, narrower fusion heights and longer fusion lengths enhance the improvement effect. The tip winglet provides the most significant effect, reducing total pressure loss by 0.64% and leakage by 16.8% when the fusion height is 1% blade height and the length is the whole chord.