Total Pressure Loss Research Articles

Experimental study on aerodynamic performance of turbine blade squealer tip with different trailing edge designs under high-speed relative casing motion

Over-tip leakage (OTL) flow leads to great aerodynamic performance reduction in high-pressure turbine, reasonable blade tip design can effectively control the loss caused by OTL flow. The relative casing motion is one of the key boundary conditions that can significantly influence OTL flow. In this study, aerodynamical tests were conducted for a high-pressure squealer tip with different trailing edge designs of full cavity squealer tip, pressure-side cutback and suction-side cutback at both stationary and rotating conditions. Loss distribution and blade near tip loading of different trailing edge structures at high-speed rotating condition are firstly reported and evaluated. The result indicates that pressure-side cutback design significantly increases the aerodynamic loss compared with full cavity tip, while suction-side cutback design has close overall loss to full cavity tip. This conclusion was consistent with numerical simulation based on Reynolds-averaged Navier–Stokes computational fluid dynamics (CFD) analysis, which reveals that pressure-side cutback design causes the cavity vortex leaks earlier compared with full cavity tip, a small vortex formed at cut region forms and results in higher total pressure loss. The effect of relative motion between blade tip and shroud reinforces this trend. The total pressure loss difference can reach 42 % compared with 8 % in stationary condition, which indicates that relative casing motion needs to be take into consideration when ranking different tip sealing designs.

Role of different cavity flame holders on the performance characteristics of supersonic combustor

Abstract The study’s primary goal was to assess the effects of different geometrical configurations of cavity flame holders on the performance characteristics of an H2-based scramjet combustor overall efficiency, focusing on mixing efficiency, combustion performance, and total pressure loss. Key findings indicate that incorporating a cavity into the combustor design leads to developing a robust recirculation zone within the cavity area. This recirculation zone is pivotal in enhancing fuel-air mixing and combustion efficiency, with cavity-based combustors showing an earlier onset of combustion and achieving peak combustion efficiencies around 90–95 %. The extent of the recirculation region is notably influenced by the proximity of the strut injector to the cavity’s length. This shock interaction can adversely affect scramjet combustor performance, especially at higher cavity angles (α) and specific geometric configurations, such as an L/D ratio of 4 and α = 30°. This advantage is critical, especially considering the compact length of the combustor, which is a desirable attribute in scramjet design.

Low-order-model-based design of additive-manufactured heat exchanger

Abstract This paper discusses a comprehensive approach to the design of an additive-manufactured heat exchanger for a high-temperature application. This manufacturing route is attractive because of the flexibility afforded in design, but there is a requirement for calibration of internal friction factors and heat transfer coefficients because of uncertainty regarding the characteristics of the manufactured internal surfaces. The design was informed by physics-based low-order models (total pressure loss model; conjugate heat transfer model) calibrated against experimental data. To extract heat transfer coefficients we use transient experiments performed over timescales for which the surfaces can be regarded as semi-infinite. The calibrated models are used to predict accurate performance characteristics for the heat exchanger. Experimental results for a prototype design were found to be in excellent agreement with the model over a wide operational range, validating both the design methodology and the utility of the particular design for the given application. Our process, though not unique, has wide applicability, has been conducted with considerable care, and will be of interest to those engaged in low-order-model-based heat exchanger design.

Positional Effects of a Fly’s Wing Vein in the Asymmetric Distribution of Hydraulic Resistances

Insect wing vein networks facilitate blood transport with unknown haemodynamic effects on their structures. Fruit flies have the posterior cross vein (PCV) that disrupts the symmetry of the network topology and reduces the total pressure loss during blood transport; however, the impact of its various positions among species has not been examined. This study investigated the haemodynamic effects of this vein with various connecting positions. By analogising venous networks to hydraulic circuits, the flow rates and pressure losses within the veins were derived using Poiseuille’s and Kirchhoff’s laws. The results showed that the total pressure loss decreased for both PCV connections near the wing’s base. In an idealised circuit imitating the network topology, applied high hydraulic resistances as one-sided as those along the edge of the wing, the same pressure loss response as that in the actual network was demonstrated, but not within a symmetric resistance distribution. Therefore, the most proximal PCV minimises the pressure loss within the asymmetric resistance distribution, indicating an evolutionary adaptation to reducing the pressure loss in certain species with this vein near the base. Our findings highlight the possible optimisation of the flies’ wing morphology to maintain the functions of the liquid transport networks and flight devices simultaneously.

Flow field explorations and design improvements of a hybrid rocket motor LOx feed line

The oxidizer system in a hybrid rocket motor needs to deliver the flow from a pressurized storage tank to multiple combustor ports. Pressure losses in the oxidizer system directly impacts combustor pressure and consequently the vehicle performance. However, oxidizer feed line designs till date have been done using simple 1D tools. Higher fidelity flow analysis methods have not been reported in the literature to identify loss generating features. Therefore, a design improvement study was carried out to identify and alleviate the impact of undesirable flow features in a typical oxidizer system design. An experimentally calibrated 3D RANS approach is applied to a typical LOx feed system which includes steps, splitters, ports, and pipes with multiple bends. These design features result in varying degrees of flow separation, secondary flows and vortical flow features and result in total pressure losses of up to 7 %. This loss means that the storage tank needs to be pressurized further to accommodate such losses and ensure combustor performance. A targeted design improvement approach that features simple, alternative, implementable solutions in the loss-generating regions is discussed. The best of these design improvements can reduce the total pressure loss to 4 %, indicating a 43 % reduction in the losses and reduced impact on storage tank design and combustor performance. Therefore, this paper demonstrates that a higher fidelity design enhancement process of the oxidizer feed system, which is often neglected in such detailed studies, can result in overall vehicle level design improvements to ensure mission targets are met effectively.

Numerical investigation of flow and heat transfer on turbine guide vane leading edge slot film cooling

As the demand for enhanced turbine cooling performance grows, the optimization of traditional discrete film hole is becoming increasingly complex. Compared with discrete holes, slot cooling is considered one of the most effective forms of film cooling, which effectively improves cooling performance while retaining a relatively simple structure. In order to explore the flow and heat transfer properties of leading-edge slot film cooling in turbine vanes, this paper analyzes the AGTB guide vane profile to examine the properties of the slot structure under different slot pitch-to-width ratios and blowing ratios (M), deeply investigates the evolution and underlying mechanism of the secondary flow downstream the slot and within the vane passage, comparisons are also made with standard cylindrical hole and inclined hole. The results show that the development of the counter-rotating vortex pairs on the pressure surface lags and extends over a longer distance downstream of the slot. Reducing the slot spacing can enhance the mutual interference between the vortex pairs, significantly improving the cooling performance on the pressure surface. On the suction side, the coolant streams remain largely independent, and both increasing M and reducing slot spacing adversely affect cooling performance. Under the same M, the slot structure achieves an average cooling effectiveness 2–3 times higher than the hole structure, accompanied by a 25 % increase in total pressure loss. At the same coolant mass flow, the cooling effectiveness on the pressure surface increases by approximately 30 %, and on the suction surface, it is three times greater than that of the hole structure, while the total pressure loss remains similar.

Study on the influence of labyrinth seal structure on the performance of shrouded stator cascades with inter-stage cavities

Numerical simulations of a shrouded stator cascade with an inter-stage cavity were conducted using validated numerical methods to investigate the impact of different labyrinth seal structures and positions on the aerodynamic performance of the stator cascade. The study discussed the development trends of secondary flow motion and corner separation within the cascade under different labyrinth seal structures, evaluated the stator performance using total pressure loss coefficient and entropy-based loss coefficient, and revealed the interaction between cavity leakage flow and mainstream flow. The results indicate that individually changing the labyrinth seal structure or position has a negligible impact on stator performance, but simultaneously altering both the structure and position of the labyrinth seal can effectively reduce stator losses. The structure case 1-position case 2 labyrinth seal structure can reduce total pressure loss by 6.25%, while structure case 4-position case 4 reduces entropy increase loss coefficient by 7.67%. Additionally, it was found that changing the labyrinth seal structure and position can increase the circumferential velocity and momentum of cavity leakage flow at the cavity outlet, enhance the resistance of cavity leakage flow to circumferential secondary flow within the stator passage, and reduce stator losses. Altering the structure and position of the labyrinth seal effectively reduces the cavity leakage vortex at the leading edge of the cascade during the mixing process of cavity leakage flow with the mainstream, causing the passage vortex to move rearward axially.

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.