Total Pressure Loss Research Articles

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%.

Experimental study on flow field and performance of bionic-wavy leading edge in an axial compressor with positive bowed blades

To enhance the aerodynamic performance of compressors in advanced aeroengines, a compound flow control method combining positively bowed blades and bionic-wavy leading edges is proposed for improving the aerodynamic performance of compressor cascades with controlled diffusion airfoils. This study verified the effectiveness of the compound control method through low-speed wind tunnel experiments using five-hole probe measurements and surface oil-flow visualization techniques. Additionally, the flow field structure was analyzed, and vortex models were established to thoroughly discuss the mechanism of the compound flow control method. The results show that within the incidence angle range of 0°–4° studied in this paper, the composite control method achieved significantly effective control, with a maximum reduction in overall total pressure loss of 25.8% compared to straight blade cascades. Three vortex models were established. The positive bowed blade cascade induced a complex vortex structure in the concentrated shedding vortex region, increasing losses in the concentrated shedding vortex (CSV) region but reducing profile losses. The coupled method further reduced profile losses and optimized the flow field in the CSV region. This study not only validates the feasibility of the compound method but also provides guidance for applying flow control methods to bowed blade cascades.

Analysis of film cooling and flow resistance characteristics of turbine blades with thermal barrier coatings

This paper employs numerical simulation to investigate the influence of thermal barrier coatings' (TBCs) thickness and surface roughness on the cooling and flow resistance characteristics of turbine blades. The results indicate that the application of TBCs significantly enhances the surface cooling efficiency of the blades. Turbine blades coated with a 0.4 mm thickness of TBC compared to blades without thermal barrier coatings, the average cooling efficiency of the blade surface increases by 12.3 %, and the maximum temperature drop at the leading edge(LE) is 317.8 K. However, the small increment in TBCs thickness leads to an increase in aerodynamic losses in the vane passage. The static pressure coefficient continuously decreases in the suction side(SS) within the interval 0.3 < x/C < 1.0. The average flow coefficient of the film holes exhibits distinct variations in different regions of the blade. As surface roughness increases, the cooling efficiency on the SS and pressure side(PS) of the blade decreases, while the heat transfer enhancement at the LE and cooling efficiency improve. Compared to a smooth coated surface, when the surface roughness height increases to 20 μm, the blade surface cooling efficiency decreases by 1.12 %, and the average temperature rise is 10.3 K. Simultaneously, the energy loss coefficient and total pressure loss coefficient in the vane passage rise with the increase in surface roughness, while the variation in the average flow coefficient of the film cooling holes remains relatively small.

Investigation of aerodynamic loss mechanism on turbine leading edge with slot film cooling based on entropy analysis

Discrete film holes are a widely used cooling structure, with increasing cooling performance demands making their optimization increasingly complex. The structure of slot cooling is simpler compared to discrete holes, but it results in greater flow losses while improving cooling performance. To elucidate the sources and mechanisms of losses in slot cooling, this study employs numerical simulations based on the Reynolds-averaged Navier-Stokes (RANS) method. The investigation focuses on the leading-edge cooling of high-pressure turbine guide vanes, using entropy analysis to thoroughly examine the effects of blowing ratio (M) and slot spacing on loss location, composition, characteristics, and mechanisms. The results indicate that the losses associated with hole structures are not sensitive to changes in blowing ratio, while the entropy generation in slot structures generally increases with higher blowing ratios. Near the corner region, the traction effect of the horseshoe vortex expands the thermal loss region of the coolant, weakening the heat exchange in hole structures but significantly enhancing it in slot structures. The hole/slot structures have a minimal impact on the location of the passage vortex core, reducing the cross-flow velocity and cross-flow layer thickness. The greater flow losses are primarily due to increased passage vortex intensity. Slot cooling can achieve superior cooling performance at low M compared to hole structures at high M, with both having similar total pressure loss. However, the thermal losses for slot structures increase by approximately 55 % compared to hole structures on the pressure side of the leading edge, with even more significant increases on the suction side, while the maximum increase in viscous losses is only about 10 %.

Toward high-lift low-solidity design for incidence tolerant gas turbine blade profile

Gas turbines operating under wide operational conditions require minimizing losses across a wide range of incidences. This study employed a blade profile efficient within a 50° incidence range to assess the effect of solidity on performance across wide incidences. The aim was to evaluate the feasibility of adopting high-lift, low-solidity designs for incidence-tolerant blade profiles. The results showed that solidity has both positive and negative effects on loss across a wide range of incidence angles. Increasing the pitch-to-axis chord ratio (the inverse of solidity) from 0.8 to 1.1 not only reduces the blade count by 37.5 % but also enhances performance at negative incidences, reducing the total pressure loss coefficient by up to 17.5 %. However, at positive incidences, the losses increased significantly as the solidity decreased. To expand the operational range of the low-solidity blade profile at positive incidences, strategic integration of leading-edge refinement and maximum deflection adjustment was applied to precisely control losses. This enabled the low-solidity blade profile to maintain lower loss levels compared to conventional high-solidity profiles up to an incidence of 7°, at which point the losses become comparable. The findings indicate that employing high-lift designs with optimized leading-edge and maximum deflection for incidence-tolerant blade profiles is feasible.

Investigation on the effect of the leading angle of the strut on the stabilization characteristics of lifted jet flames

To achieve stable combustion within supersonic combustors, the strut-based mixing augmentation techniques are commonly employed. Large Eddy Simulations (LES) are employed to capture the intricate details of the combustion process in this study, focusing on the influence of shock waves induced by the leading edge of the strut and recirculation zones formed at the trailing edge on flame structure and stability. The effects of five struts with different leading edge angles on the stabilization characteristics of lifted jet flames were investigated, and a novel wedge-shaped strut was proposed. The results reveal that variations on the leading-edge angle of the struts lead to significant differences in the morphology of the jet flames. Under conditions permitting stable combustion, a strut with a 5° leading edge half-angle exhibits the highest combustion efficiency, though the combined effects of shock waves and recirculation zones in the 7° leading edge half-angle strut are also noteworthy. Based on these findings, a novel wedge-shaped strut design is introduced, which combines the characteristics of both strut types. The novel wedge-shaped strut achieves a combustion efficiency of 97 % at the slight cost of additional total pressure loss, providing a foundation for further optimization.

Numerical Study on the Influence of Inlet Turbulence Intensity on Turbine Cascades

The turbulence intensity of the high-pressure turbine inlet plays an important role in the development of secondary flow and boundary layer evolution in the turbine passage. Unfortunately, current research often overlooks the coupling effect between boundary layer separation and endwall secondary flow, and lacks comprehensive exploration of loss variation. To complement existing research, this study utilizes numerical simulation techniques to investigate the evolution of the boundary layer and secondary flow in high-pressure turbine cascades under varying turbulence intensities, with experimental research results as the basis. Furthermore, the relationship between profile loss and endwall secondary flow loss is analyzed. The research results indicate that higher turbulence intensity can enhance the anti-separation ability of the boundary layer, thereby reducing the blade profile loss caused by separation in the boundary layer. However, higher turbulence intensities (Tu) enhance the development of secondary flow within the cascade, leading to a significant increase in secondary flow losses. Q-criterion methods are employed to display the vortex structures within the passage, while loss decomposition is leveraged to uncover the variations of different losses under different turbulence intensities (Tu of 1% and 6%).With the exception of a minor decrease in total pressure loss (Yp) when Tu increases from 1% to 2%, Yp demonstrates a substantial increase in all other cases with increasing Tu. Additionally, this study explains why the loss of high-pressure turbine cascades shows a trend of first increasing and then decreasing with the increase in inlet turbulence intensity.

Multi-objective optimization design of aerothermal performance for turbine blade squealer tip with inclined suction-side rim

To optimize the inclined rim design, multi-objective optimization was numerically conducted on the suction-side rim of the squealer tip in the GE-E3 rotor blade. The RANS equations, incorporating with the standard k-ω model, were numerically solved. Through the implementation of a Kriging surrogate model coupled with the NSGA-II evolutionary optimization technique, six geometric parameters on the suction-side rim went through iterative optimization. This process yielded three optimized squealer tip geometries on the Pareto front: Structure 1 with minimum total pressure loss, Structure 2 with optimal comprehensive aerothermal performance, and Structure 3 with minimum tip heat load. Comparative analysis revealed that with the inner and external wall inclined toward the cavity and passage respectively, the optimization results in a 3 %–11 % reduction in tip heat load, accompanied by an 8.5 %–10.2 % reduction in aerodynamic loss, facilitated by an attenuated upper passage vortex (UPV) and strengthened tip leakage vortex (TLV). The critical inclined region of the external rim wall, spanning 7 %–80 %, influences the exit total pressure loss distribution. The inclined inner wall upstream of 50 % Cax obstructs coolant jets from exiting the cavity, reducing the heat load on the cavity bottom. However, the inclined external wall from 66 % to 90 % Cax amplifies the impingement of the scraping vortex (SV), increasing local heat transfer. The thermal behavior of Structure 1 is sensitive to changes in tip clearance, whereas Structure 2 and 3 show more robust advantages. As the clearance increased, the aerodynamic losses associated with the TLV rise for all three optimized tips, but losses from the UPV remain similar. This study provides a novel approach for designing efficient squealer tip geometries incorporating an inclined rim.

Influence of mixed dimples on aerodynamic performance of high-load compressor cascade

In this study, the effect of mixed dimples on the aerodynamic performance of a high-load compressor cascade, named National Advisory Committee for Aeronautics 65-K48, at different attack angles was numerically and experimentally investigated. Mixed dimples with four columns and eight rows were arranged at 10% to 32% of the chord length of the half-blade height. The performance parameters of the cascade outlet section and vortex structure inside the cascade were exploited. The results reveal that the boundary layer is disturbed, and the boundary layer transition occurs ahead of time, induced by the mixed dimples. The anti-separation ability of the fluid was enhanced, and the separation bubbles on the suction surface were eliminated. Meanwhile, the lateral secondary flow was inhibited, and the intensities of the passage and concentrated shedding vortices were consequently reduced. Compared with the prototype cascade, the total pressure loss of the dimpled cascade at all attack angles was reduced. The total pressure loss can be reduced by nearly 20% at –9°.

On the Challenges of Integrating a Rotating Detonation Combustor with an Industrial Gas Turbine and Important Design Considerations for Row-1 Blades

Abstract In an effort to increase the efficiency and performance of gas turbine power cycles, pressure gain combustion (PGC) has gained significant interest. Since Rotating Detonation Combustors (RDC) can provide a quasi-steady mode of operation, research has been triggered to integrate RDC with power-generating gas turbines. However, the presence of subsonic and supersonic flow fields which are generated due to the shock waves that stem from the detonation wave front and the highly non-uniform temperature and velocity profiles may cause a depreciation in the turbine performance. The current study seeks to investigate the challenges of integrating the RDC with nozzle guide vanes (NGV) of an industrial, can-annular gas turbine and attempts to understand the major contributors that impact efficiency and identify the key areas of optimization that need to be considered for maximizing performance. The RDC was integrated with the NGVs through a non-optimized straight duct-type geometry with a diffuser cone. 3-Dimensional Numerical analyses were performed to investigate sources of total pressure loss and to understand the unsteady effects of RDC which contribute towards the deterioration of performance. The entropy generation at different regions of interest was calculated to identify the major irreversibilities in the system. Also, total pressure and temperature distribution along the radial direction at the exit of the transitional duct is presented.

Experimental Study of High-Diodicity Inlet Effects On Rotating Detonation Combustor Performance and Operability

Abstract Implementing high-diodicity inlets is critical to reduce backflow and mitigate the pressure loss across the injector in RDCs. Experiments on air injection diodicity were conducted in a non-premixed RDC for both cold-flow and reacting conditions. The introduction of a Tesla-like diode impacted operating modes and injector dynamics, but the extent of that effect depended on the throat-to-combustor area ratio. A smaller ratio mitigated the impact of the diode on detonation characteristics, while a larger ratio extended the operating range of stronger wave modes. The diode stabilized RDC operation through an increased static pressure drop, but limited performance most likely due to poor reactant mixing and local equivalence ratio distribution. Cold-flow tests showed a higher diodicity for the diode, which may contribute to higher pressure gain in reacting experiments. A modified diodicity formulation based on reacting flow measurements was introduced, suggesting that a multi-metric approach can be useful to assess injector performance. High-diodicity air inlets could be useful tools for reducing total pressure loss and controlling operating modes, but careful consideration is required to limit adverse effects on processes like reactant mixing.

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.

The loss mechanisms and vortex dynamics of rotor platform step geometries in a low-speed research compressor stage

The construction and assembly of an axial compressor blade row introduces real geometric features, which leads to the deviation of flow characteristics from the intended design. According to the realistic assembly errors observed in the four-stage low-speed research compressor developed by Shanghai Jiao Tong University, two representative step geometries at the rotor platform were abstracted. Unsteady simulations were carried out in the stage environment to study the influence of hub discontinuities on compressor performance and loss mechanisms. The comparisons of smooth hub, elevated hub (EH), and a wedged hub (WH) demonstrated that all the step geometries resulted in increased loss inside both rotor and stator rows. The loss induced by WH was more sensitive to step heights than EH. Under the influence of forward facing step, rotor incidence was reduced and horseshoe vortex (HV) was exacerbated, leading to the accumulation and separation of low momentum fluids at the pressure side corner. This led to a redistribution of losses across the span. In the lower half span, the relative total pressure loss increased almost linearly with step heights, while in the tip region, it reduced slightly. The backward facing step generated the step corner vortices (SCV) that shed and propagated downstream. The SCV interacted with the HV and cavity leakage vortex in the stator passage. Coupled with the augmented stator inflow angle, the corner separation at the suction side was dramatically intensified and led to a significant loss increase.

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.

EFFECTS OF CONTINUOUS AND FRACTURED RIBS PLACED AT THE ENDWALL SURFACE UPSTREAM OF THE TURBINE-VANE LEADING-EDGE

Total pressure loss, chaotic and heavy flow downstream of the turbine passage exit area are a few effects of the desired increase in inlet temperature input at the upstream part of the turbine vane passage. Huge amount of energy from the cross-flows near the middle of the vane actual chord are transported downstream by this chaotic flow which often results in aerodynamic loss and differential heat penalties. To compare the different arrangement, comparisons of the two geometries, the impact of adding various rib configurations upstream are studied for heat augmentation efficacies. The leading-edge rib(s) effects at the mid-stream of the vane passage are quantitatively examined in this work using computational fluid dynamics (CFD) technique. With the employment of Reynolds Averaged Navier-Stoke energy equations, the geometries are modelled choosing the appropriate boundary conditions. Polyhedral mesh is used with fifteen prism layer mesh at the endwall and vane surfaces to capture the flow physics close to the endwall area. Although the two configurations employed showed relative reduction in the total pressure loss coefficient, however, the fractured ribs produced superior outcomes of Nusselt number reduction of over 10% along the passage exit region. The data demonstrate a considerable difference in the overall pressure loss between the two configurations.

The effect of including dynamic imaging derived airway wall motion in CFD simulations of respiratory airflow in patients with OSA

Obstructive sleep apnea (OSA) is an airway disease caused by periodic collapse of the airway during sleep. Imaging-based subject-specific computational fluid dynamics (CFD) simulations allow non-invasive assessment of clinically relevant metrics such as total pressure loss (TPL) in patients with OSA. However, most of such studies use static airway geometries, which neglect physiological airway motion. This study aims to quantify how much the airway moves during the respiratory cycle, and to determine how much this motion affects CFD pressure loss predictions. Motion of the airway wall was quantified using cine MRI data captured over a single respiratory cycle in three subjects with OSA. Synchronously-measured respiratory airflow was used as the flow boundary condition for all simulations. Simulations were performed for full respiratory cycles with 5 different wall boundary conditions: (1) a moving airway wall, and static airway walls at (2) peak inhalation, (3) end inhalation, (4) peak exhalation, and (5) end exhalation. Geometric analysis exposed significant local airway cross-sectional area (CSA) variability, with local CSA varying as much as 300%. The comparative CFD simulations revealed the discrepancies between dynamic and static wall simulations are subject-specific, with TPL differing by up to 400% between static and dynamic simulations. There is no consistent pattern to which static wall CFD simulations overestimate or underestimate the airway TPL. This variability underscores the complexity of accurately modeling airway physiology and the importance of considering dynamic anatomical factors to predict realistic respiratory airflow dynamics in patients with OSA.

Experiment investigation on secondary flow loss and flow control mechanisms in a variable compressor cascade with penny cavities

Variable Stator Vanes (VSVs) with penny cavities are an important method to improve the efficiency of compressors. However, the annular and radial gaps in VSVs introduce complex flow structures, posing challenges in enhancing compressor efficiency. This article employs low-speed wind tunnel experiments to study the impact mechanisms of VSVs' adjustment angle and radial gap size on secondary flow losses and implements single-hole suction as a means of flow control to reduce total pressure loss. The results indicate that radial gap leakage has a certain inhibitory effect on annular gap leakage at general adjustment angles. Total pressure loss decreases initially as the radial gap increases but once the radial gap further enlarges and dominates the loss components, the total pressure loss increases proportionally with the radial gap. At extreme adjustment angles, where the lateral pressure gradient is significant, radial gap leakage intensity is high and comprises the major part of the loss, leading to an increase in total pressure loss in proportion to radial gap enlargement. Additionally, single-hole suction is effective under various conditions, with the potential to reduce total pressure loss by up to 19.4 %. These findings provide guidance for the application of VSVs in further improving compressor efficiency.

Research on flow and heat transfer characteristics of supersonic film cooling with different hole-configurations

In order to investigates the flow and heat transfer characteristics under supersonic mainstream conditions among three typical film hole-configurations—cylinder hole (CH), fan-shaped hole (FSH), and laidback fan-shaped hole (LFSH), this study conducted numerical simulation using the ANSYS CFX and based on the assumption of compressible flow. The results indicate that the cooling effectiveness of the CH dramatically decreases as the blowing ratio increases, while FSH and LFSH improves. When the supersonic mainstream encounters the coolant, flow structures such as oblique shock waves, expansion waves, and mixing layers et al., will be generated around the film holes, which have an impact on the flow and heat transfer characteristics. Under the conditions of same blowing ratio, the shock wave intensity of LFSH is significantly lower than the other two holes and the kidney vortices formed by the CH develop further along the mainstream direction. While the strength of the kidney vortices formed by the FSH and the LFSH is weaker. Additionally, the LFSH incurred the least total pressure loss from before the shock wave to after the oblique shock wave. Specifically, at a blowing ratio of 0.8, it achieved a reduction of approximately 20% in total pressure loss compared to the CH.

Performance prediction of co-rotating disk cavity with finned vortex reducer based on machine learning

Machine learning method was applied for rapid and precise prediction of flow and heat transfer characteristics within co-rotating disk cavity with a finned vortex reducer. A new data preprocessing scheme was developed to reduce modeling cost. By using this scheme, classical radial basis function neural network (RBFNN) shows better prediction performance compared with deconvolutional neural network. Furthermore, RBFNN has a simpler topological structure and fewer hyperparameters. By testing, the relative root mean square error for total pressure, total temperature, and swirl ratio is 0.51 %, 0.32 %, and 1.99 %, and corresponding coefficient of determination reaches 99.68 %, 96.55 %, and 99.16 %. The effects of input parameters on the outputs of RBFNN were analyzed, and a sensitivity analysis was conducted. Within the investigated parameter range, the total pressure loss increases as dimensionless mass flow rate, rotational Reynolds number, and radial position of the fin increases. The total temperature drop increases with the increase of dimensionless mass flow rate and rotational Reynolds number, but the decrease of radial position of the fin. Additionally, an increase in dimensionless mass flow rate or a decrease in rotational Reynolds number causes the increase of swirl ratio. This research provides an efficient modeling method for components in secondary air system in gas turbines.

Numerical investigation on cooling performances of laminated cooling configuration with external discrete slot-trench film holes and internal clapboards

Based on a typical laminated configuration, a new cooling configuration that integrates internal and external enhancements is proposed in this paper. For external enhancement, discrete slot-trench film holes are utilized, while clapboards are introduced for internal improvement. Numerical results indicate that the extensibility and adhesiveness of the film are significantly enhanced by the discrete slot-trench film holes, resulting in an increase in the average adiabatic film cooling effectiveness of the externally improved configuration by 60.13% to 141.86% compared to the laminated configuration. The internal heat transfer area is increased by 55.31% by the clapboards, resulting in an average internal cooling effectiveness increase of 1.29% to 3.99% for the coupled improved configuration compared to the externally improved configuration. Due to these positive cooling effects, the coupled improved configuration shows an increase in average overall cooling effectiveness ranging from 11.55% to 14.39% under different blowing ratios, compared to the laminated configuration. Furthermore, the performance evaluation criteria for the coupled improved configuration are all greater than 1.0 under test conditions, indicating that the enhancement of overall cooling effectiveness does not cause significant total pressure losses.