Axial Chord Length Research Articles

The Influence of Bleed Position on the Stability Expansion Effect of Self-Circulating Casing Treatment

The self-circulating casing treatment can effectively expand the stable working range of the compressor, with little impact on its efficiency. With a single-stage transonic axial flow compressor NASA (National Aeronautics and Space Administration) Stage 35 as the research object, a multi-channel unsteady numerical calculation method was used here to design three types of self-circulating casing treatment structures: 20% Ca (axial chord length of the rotor blade tip), 60% Ca, and 178% Ca (at this time, the bleed position is at the stator channel casing) from the leading edge of the blade tip. The effects of these three bleed positions on the self-circulating stability expansion effect and compressor performance were studied separately. The calculation results indicate that the further the bleed position is from the leading edge of the blade tip, the weaker the expansion ability of the self-circulating casing treatment, and the greater the negative impact on the peak efficiency and design point efficiency of the compressor. This is because the air inlet of the self-circulating casing with an air intake position of 20% Ca is located directly above the core area of the rotor blade top blockage, which can more effectively extract low-energy fluid from the blockage area. Compared to the other two bleed positions, it has the greatest inhibitory effect on the leakage vortex in the rotor blade tip gap and has the strongest ability to improve the blockage at the rotor blade tip. Therefore, 20% Ca from the leading edge of the blade tip has the strongest stability expansion ability, achieving a stall margin improvement of 11.28%.

Enhanced turbine endwall cooling using crescent-dimpled film cooling holes and a hybrid configuration: numerical study

The current study proposed a novel crescent-dimpled film cooling hole and investigated its aerothermal performance on the turbine endwall when a row of the film holes are arranged in front of the stator. The mainstream Reynolds number based on the inlet velocity and the axial chord length of the vane was 150,000, and the blowing ratio ranged from 0.5 to 1.5. Besides, a hybrid configuration combining the advantages of the cylindrical and crescent-dimpled holes was also investigated. RANS simulations using Shear Stress Transport (SST) k–ω turbulence model were conducted. The numerical simulations show that the pure crescent-dimpled hole design enhances the adiabatic film cooling effectiveness by 32.3 %, 52.3 %, and 43.6 % at blowing ratios of 0.5, 1.0, and 1.5, respectively. Correspondingly, the net heat flux reduction (NHFR) values are 22.9 %, 57.9 %, and 63 % higher than the cylindrical holes. The high film cooling effectiveness behind the film cooling hole prevents the additional thermal load caused by the dimple-induced heat transfer enhancement. Using cylindrical holes near the leading edge and crescent-dimpled holes elsewhere, the hybrid arrangement suppresses the passage vortex and further enhances the film cooling effectiveness and NHFR by 51.9 % and 93.8 % at BR = 1.5, respectively. The streamlines and vortex structures show that the crescent dimple at the hole’s exit diffuses the coolant, thereby enhancing the film cooling in the lateral direction. Flow separation occurs behind the dimple, which reduces the jet momentum and attracts the jet towards the wall. The curved surface of the dimple directs the horseshoe vortex in front of the jet to the side, and anti-CRVP is formed. These are responsible for the film cooling enhancement by the crescent dimple.

Exploring the Aerodynamic Effect of Blade Gap Size via a Transient Simulation of a Four-Stage Turbine

With the impact of size on low-pressure turbines (LPTs) increasing, the gap between the blades has shrunk, inevitably influencing the unsteady effects inside the turbine. In this study, the aerodynamic effect of the blade gap size is investigated using a compressible unsteady Reynolds-averaged Navier–Stokes (URANS) model on the basis of a four-stage LPT. Simulations are conducted in which the gap between the third-stage stator (S3) and rotor (R3) varies from 0.2 to 0.8 times the axial chord length of the R3 blade. The multi-stage environment reflects the complexity of real low-Reynolds flow fields. Computational fluid dynamics is used to analyze the flow field in detail. The results demonstrate that in the small-gap (AG-0.2) case, the turbulence kinetic energy (TKE) level of the S3 wake close to the R3 leading edge is four-thirds of that in the large-gap (AG-0.8) case. The higher intensity of the wake impacting on the blade results in a higher inverse pressure gradient in the rear part of the R3 suction surface, which increases the profile loss. However, the AG-0.2 case leads to fewer losses caused by the passage vortex in the hub area under the influence of the higher intensity of the wake.

Control of separation flows of turbine-blade-tip turbines by splitter blades

Gas turbines cannot be directly reversed, and the astern operation of gas turbines still requires additional power transmission equipment to be achieved. To compensate for this deficiency, the application of turbine-blade-tip turbine in gas turbines was proposed. Turbine-blade-tip turbines have the characteristics of extremely low solidity and extremely high diameter-span ratio. Compared with conventional turbine blades, the capacity of turbine blades to restrain gas flow is greatly weakened, the separation flow is obvious when going astern, resulting in low efficiency. For the control problem of low solidity turbine-blade-tip turbines separation flow, a splitter blade control strategy is selected to suppress the separation flow. Adding a splitter blade between original rotor blades effectively improves the ability of rotor blades to restrain gas flow, significantly reduces separation phenomena, and significantly improves efficiency. As the axial chord length of the splitter blade shortens, the inhibitory effect on the separation flow decreases, and the ability of splitter blades to improve the efficiency gradually descends.

Influence of bleeding positions in self-recirculating casing treatments on the stability of a subsonic axial compressor

Influence of bleeding positions in self-recirculating casing treatments on the stability of a subsonic axial compressor

Film cooling effectiveness from upstream purge slots of a turbine vane endwall: Experiment, modeling, and correlation

The design of an efficient cavity slot between assembled components, where purge air is injected, is especially crucial to the durability and efficiency of an aero-engine, since the purge air is relevant to sealing and film cooling of the hot component. In this paper, film cooling for purge flow injected from slots upstream the endwall of a turbine vane is systematically investigated in a linear cascade using Pressure-Sensitive Paint (PSP). The influence of the slot width (0.020−0.044 times vane axial chord length), position (0.11−0.24 times vane axial chord length in front of turbine passage inlets), and inclination angle (20−45°) is examined for coolant injection ratios of 0.3−2.0 %, with an attempt to develop a new correlation specifically for purge flow film cooling over a turbine endwall, since existing correlations in literature only work for slot injection on flat plates. The experimental tests are undertaken in extensive ranges of the slot geometry and the injection rate that are typically seen in a modern engine. For better understanding the impacts of the affecting parameters, five-hole probe and thermocouple measurements and numerical simulations were carried out as well to visualize the coolant-with-mainstream mixing flows. The measured local film coverage, pitchwise- and area-averaged cooling effectiveness are presented and discussed, in combination with the probe-detected and numerically-modeled flow structures to elaborate the experimental observations. The current results sufficiently reveal the changing tendency of purge flow film cooling caused by the variations of the slot geometry. Additionally, the newly-developed correlation is well validated using own experimental data and datasets in literature, proving the capability of the correlation to predict film cooling for purge flow injected from slots with such geometrical parameter ranges. The correlation can be a helpful tool to support the selection of an efficient slot in its preliminary design stage.

Large eddy simulations and experiments on low-Reynolds-number laminar separation control for low-pressure turbine blades with oblong-dimpled surface

Modern high-performance low-pressure turbines (LPT) usually encounter flow problems at low Reynolds numbers, especially for high-altitude UAVs under cruise conditions. Flow separation may occur on the blade suction side, seriously degrading the turbine's aerodynamic efficiency. The present study investigates a passive laminar flow separation control method based on the oblong dimple surface structures for their more potent flow acceleration abilities. Linear cascade tests and large eddy simulations are conducted for a typical LPT blade profile, T106A, at the Reynolds numbers of 25,000, 50,000, and 100,000 based on the inlet velocity and axial chord length. Results show that applying the oblong dimples can apparently reduce the reverse flow region and improve the aerodynamic performance at low Reynolds numbers. The experimental results achieve a maximum total pressure loss reduction of 34.7% at a low Reynolds number of 25,000. The numerical simulations show the local flow acceleration after the oblong dimples and demonstrate two working modes. At low Reynolds numbers, the oblong dimples accelerate the fluid above, generating local high-speed flows, and the uneven velocity distributions in the spanwise direction feed the deformation and fragmentation of the downstream separating vortex. While at high Reynolds numbers, the oblong dimples act as vortex generators, leading to flow transition on the blade suction surface.

Film Cooling Effectiveness on Pressure Surface and Suction Surface of Turbine Guide Vane With Diffusion Slot Holes

Abstract This paper investigated the film-cooling effectiveness of diffusion slot holes in a turbine nozzle guide vane. The pressure-sensitive paint measurement technique was employed to obtain the film-cooling effectiveness at a density ratio of DR = 1.5. The mainstream Reynolds number based on the axial chord length and the exit velocity was 6 × 105. The mainstream turbulence intensity was approximately 3.7%. Three diffusion slot hole geometries with cross-sectional aspect ratios (ASs) of 2.3, 3.4, and 4.9 were tested and compared with a typical fan-shaped hole. The experiments were performed at three typical hole row locations on the pressure surface (PS) and suction surface (SS). The average blowing ratios varied from M = 0.5 to 2.5. The results showed that throughout the blowing ratio range, on the PS, a substantially higher film-cooling effectiveness than the fan-shaped hole is always obtained from the diffusion slot hole with a large aspect ratio (AS = 4.9); on the SS, the diffusion slot hole with a small AS (AS = 2.3). The influence of hole row positioning is inconsistent for diffusion slot holes with different ASs. The diffusion slot hole is less affected by the PS when the AS is moderate and less affected by the SS when the AS is large. The film-cooling effectiveness of the diffusion slot holes is basically the lowest where the PS has a maximum concave curvature and the highest where the SS has a large favorable pressure gradient.

Investigation on Shallow Trenched-Hole Film Cooling Effectiveness at Turbine Blade Pressure Side

Numerical simulations are performed to investigate the effects of shallow trenches on the adiabatic film cooling effectiveness (ηad) of single row film cooling on the pressure side of a turbine guide vane, under the cascade representative conditions. The mainstream inlet Reynolds number is set as 4×105. A single row of film cooling holes are located at 30% axial chord length at the pressure side. Three shallow trenches with different widths (w/d=2, 2.5 and 3) but have the same depth (0.75d) and a cylindrical film cooling hole are compared in the current study at M= 2.0. The results show that the shallow trenched-hole on the pressure side significantly suppresses lift-off of the secondary flow, and effectively improves the spanwise uniformity of film coverage. Therefore, compared with the conventional cylindrical hole (CH), the shallow trenched-hole shows a significant enhanced film cooling performance at M=2.0. By comparisons, it is confirmed that the shallow trenches with w/d=2 and 2.5 are more pronounced than the shallow trench with w/d=3. While the shallow trench with w/d=2.5 takes a little advantage over the w/d=2.

Impact of the ending position of controllable speed casing on the flow stability in a transonic compressor rotor

The structure of casing treatment has a significant influence on the stability expansion effect. The ending position of the rotatable ring is one of the important structural parameters in the controllable speed casing. The impact of ending position of the rotatable ring rotating at the same speed as the rotor on the stability expansion and stability expansion mechanism was studied in this paper. The results show that, when the ending position of rotatable ring expands downstream in the range of 20% to 60% axial chord length, because of the improvement in the tip leakage vortex trajectory and main flow/leakage flow interface, the stability expansion effect rises gradually. The further downstream expansion of ending position from 60% axial chord length, however, has a slight impact on tip flow, bringing out a steady stability expansion effect around 52.5%. The best stability expansion effect of ending position occurs at 60% axial chord length where the stable operating margin increases by 53.5%.

Mechanism Affecting the Performance and Stability of a Centrifugal Impeller by Changing Bleeding Positions of Self-Recirculating Casing Treatment

This study aimed to investigate the influence of the bleeding position of a self-circulating casing on the aerodynamic performance of a transonic centrifugal compressor. Three types of self-circulating structures with the bleeding positions of 11% Ca (the axial chord length of the blade tip), 14% Ca and 20% Ca from the leading edge of the blade were studied by using the numerical simulation method, with the Krain impeller taken as the research object. It was found that all three types of self-recirculating casing treatments can expand the stable operating range of the impeller, and that at medium and small flow rates, the total pressure ratio and efficiency of the impeller increase gradually with the backward movement of the bleeding position. The self-circulating casing treatment can restrain the development of tip leakage vortex, reduce the blockage area, and improve the stability of the impeller by sucking low-energy fluid. The farther back the bleeding position is, the greater the bleeding mass flow rate of the self-circulating casing for the low-energy fluid in the blade-tip passage becomes. Additionally, a greater inhibition effect on the tip leakage vortex, and a better effect of improving the performance and stability of the impeller, can be obtained. The best air bleeding position is 20% Ca, but it is not directly above the blade-tip blockage center of the solid wall casing passage. Instead, it is downstream of the blockage area.

Augmented cooling performance in gas turbine blade tip with slot cooling

The blade tip is one of the most susceptible parts of gas turbine blade due to the thermally induced failure caused by a hot gas tip leakage flow through tip clearance. A highly sophisticated and efficient blade tip cooling scheme plays a crucial role in reliable operation of a gas turbine blade. In the present study, we adopted a special slot cooling scheme in turbine blade tip. The film cooling effectiveness (FCE) and flow characteristics of turbine blade tip were investigated according to the cooling flow through the slot gap in blade tip rim. A linear cascade was fabricated to measure and evaluate the cooling performance of the blade tip with the film cooling through a slotted rim versus the blade tip with cooling holes. The local FCE values of the blade tip floor and rim upper surface were measured using the pressure-sensitive paint (PSP) method. The Reynolds number based on the axial chord length and the inlet velocity in the experiment was set at 192,000. The tip clearance was varied from 1 to 3% of the blade span and the coolant mass flow rate ranged from 0.3 to 0.7% of the main flow. The slot cooling in the blade tip, which can be incorporated through an additive manufacturing, showed highly enhanced cooling performance ranging from 53 to 116% compared to that of the blade tip with film cooling holes for various tip gaps and coolant mass flow rates. The slot cooling in the blade tip is expected to improve the thermal durability and reliability of blade tip with better coolant coverage and additional convective cooling through the internal cooling passage in the narrow squealer rim.

Combined control of variant clearance and casing aspiration on tip flow field and instability of a transonic compressor

To enhance the transonic compressor performance, the axially non-uniform variant tip clearance schemes and combined schemes with casing aspiration were applied and their effects on the tip flow field and flow instability were studied numerically. The results show that the expanding tip clearance can increase the stall margin and peak efficiency by reducing the tip leakage flow intensity within the first half chord, and cause the interface between the leakage flow and mainstream to move downstream. The effect of the shrinking clearance is on the contrary. The compressor performance can be further improved by the combination of expanding clearance and casing aspiration. When the circumferential suction groove is located at 35% of the axial chord length corresponding to the low-velocity region due to the leakage vortex breakdown, the stability enhancement is the largest, which is about 7.2% higher than that under the uniform clearance scheme. Casing aspiration decreases the blade tip loading and attenuates the tip leakage vortex, tip secondary vortex, and their interaction, which results in the suppression of vortex breakdown and tip flow blockage. Moreover, the oscillations and instability of the flow field are weakened. Therefore, the tip flow field is modified and the compressor performance is improved.

Numerical investigation of an aggressive s-shaped compressor transition duct with boundary layer suction

Aggressive s-shaped compressor transition ducts are important components in the connection between upstream boosters and downstream high-pressure compressors. The flow path is an s-shaped channel with struts and a large radial drop length ratio, which breaks through the limitations of traditional design and has a large aerodynamic loss. Therefore, this paper considers an aggressive s-shaped compressor transition duct in a geared turbofan engine and creatively proposes a method for controlling the flow separation through boundary layer suction. The results show that hub suction reduces the losses of the aggressive s-shape transition duct. As the mass flow rate of hub suction increases, the total pressure loss coefficient decreases and the rate of reduction in the total pressure loss slows down. Combined boundary layer suction reduces the total pressure loss to a greater extent. On the premise that the location of blade suction remains unchanged, the optimal location for the circumferential slot of hub suction is at 20% of the axial chord length of the strut, whereby the total pressure loss coefficient decreases by about 30% compared with the case of no suction. When the mass flow rate of suction is fixed at 3% of the inlet mass flow rate, a distribution of 0.5% from blade suction and 2.5% from hub suction reduces the total pressure loss by 1.6% compared with the case where all 3% comes from hub suction. The distribution of the mass flow rate for combined boundary layer suction has an optimal ratio.

Large-Eddy Simulation of the Boundary Layer Development in a Low-Pressure Turbine Cascade With Passive Flow Control

The boundary layer development on a low-pressure turbine blade surface modified by recessed dimples, U-grooves, and rectangular grooves has been investigated through the large-eddy simulations. The simulations are performed at a Reynolds number of 50,000 (based on the inlet velocity and axial chord length) and extremely low freestream turbulence conditions. The characteristic parameters of the boundary layer are used to estimate the development of the boundary layer, and spectral analysis has also been performed to identify the dominant frequency of shedding vortices. The results of simulations indicate that three surface modifications all reduce the profile losses by restraining the separation bubble size. However, the grooves and dimples show different mechanisms in inhibiting laminar separation. Grooves tend to promote the formation of spanwise vortices, which is more difficult to break into turbulence. A high-speed shedding vortex is induced by the particular 3D structure of dimple, and its shedding frequency is nearly twice the Kelvin–Helmholtz (K-H) instability frequency. The interaction between the shedding vortices and the K-H vortices promotes the breakdown process of the spanwise vortices, which leads to an earlier transition of the boundary layer at a low disturbance level. The current study reveals the different mechanisms of dimples and grooves and shows the great potential of dimples for flow control in low-pressure turbines. Besides, the flow structures inside the dimples with adverse pressure gradients are also explored.

Numerical Investigation of Single-Row Double-Jet Film Cooling of a Turbine Guide Vane under High-Temperature and High-Pressure Conditions

Single-row double-jet film cooling (DJFC) of a turbine guide vane is numerically investigated in the present study, under a realistic aero-thermal condition. The double-jet units are positioned at specific locations, with 57% axial chord length (Cx) on the suction side or 28% Cx on the pressure side with respect to the leading edge of the guide vane. Three spanwise spacings (Z) in double-jet unit (Z = 0, 0.5d, and 1.0d, here d is the film hole diameter) and four spanwise injection angles (β = 11°, 17°, 23°, and 29°) are considered in the layout design of double jets. The results show that the layout of double jets affects the coupling of adjacent jets and thus subsequently changes the jet-in-crossflow dynamics. Relative to the spanwise injection angle, the spanwise spacing in a double-jet unit is a more important geometric parameter that affects the jet-in-crossflow dynamics in the downstream flowfield. With the increase in the spanwise injection angle and spanwise spacing in the double-jet unit, the film cooling effectiveness is generally improved. On the suction surface, DJFC does not show any benefit on film cooling improvement under smaller blowing ratios. Only under larger blowing ratios does its positive potential for film cooling enhancement start to show. Compared to the suction surface, the positive potential of the DJFC on enhancing film cooling effectiveness behaves more obviously on the pressure surface. In particular, under large blowing ratios, the DJFC plays dual roles in suppressing jet detachment and broadening the coolant jet spread in a spanwise direction. With regard to the DJFC on the suction surface, its main role in film cooling enhancement relies on the improvement of the spanwise film layer coverage on the film-cooled surface.

Endwall Heat Transfer and Cooling Performance of a Transonic Turbine Vane With Upstream Injection Flow

Abstract Flow fields near the turbine vane endwall region are very complicated due to the presence of high three-dimensional passage vortices and endwall secondary flows. This makes it challenging for the endwall to be effectively cooled by using traditional endwall cooling methods, such as impingement cooling combined with local film cooling inside the vane passage. One effective endwall cooling scheme: coolant injection flow through discrete holes upstream of the vane leading edge on the endwall, has been considered by many gas turbine companies. The present paper focuses on endwall film cooling effectiveness evaluation with upstream coolant injection through discrete holes. Detailed experimental and numerical studies on endwall heat transfer and cooling performance with coolant injection flow through upstream discrete holes are presented in this paper. High-resolution heat transfer coefficient (HTC) and adiabatic film cooling effectiveness values were measured using a transient infrared thermography technique on an axisymmetric contoured endwall. The endwall tested was a scaled up inner endwall of an industrial transonic turbine vane with double-row discrete cylindrical film cooling holes located 0.39Cx upstream of the vane leading edge. The tests were performed in a transonic linear cascade blowdown wind tunnel facility. Conditions were representative of a land-based power generation turbine with exit Mach number of 0.85 corresponding to exit Reynolds number of 1.5 × 106, based on exit condition and axial chord length. A high turbulence level of 16% with an integral length scale of 3.6%P was generated using inlet turbulence grid to reproduce the typical turbulence conditions in real turbine. Low temperature air was used to simulate the typical coolant-to-mainstream condition by controlling two parameters of the upstream coolant injection flow: mass flowrate to determine the coolant-to-mainstream blowing ratio (BR = 2.5, 3.5), and gas temperature to determine the density ratio (DR = 1.2). To highlight the interactions between the upstream coolant flow and the passage secondary flow combined with the influence on the endwall heat transfer and cooling performance, a comparison of CFD predictions with experimental results was performed by solving steady-state Reynolds-averaged Navier–Stokes (RANS) using the commercial CFD solver ansys fluent V.15. A detailed numerical method validation was performed for four different Reynolds-averaged turbulence models. The realizable κ−ε model was validated to be suitable to obtain reliable numerical solution. The influences of a wide range of coolant-to-mainstream blowing ratios (BR = 1.0, 1.5, 1.9, 2.5, 3.0, 3.5) were numerically studied. Complex interactions between coolant injections and secondary flows in vane passage were presented and discussed. An optimal value of the blowing ratio for the present upstream discrete film hole is also suggested based on the current study. Results indicate that for lower values of BR, the endwall coolant coverage from the upstream double-row discrete holes is strongly controlled by the passage secondary flow, thus the cooling effectiveness is very poor. As the BR increases, the strong secondary flow in vane passage can be suppressed by the coolant injections and begin to be almost eliminated when BR increases to a critical value (BR = 2.5–3.0). Beyond the critical BR, most of the injected coolant begins to lift off from the endwall and penetrate significantly into the mainstream flow, yielding inefficient endwall cooling performance.

Influence of Surface Roughness on a Highly Loaded Axial Compressor Stage Performance at Low Reynolds Number

In the present study, a numerical simulation was conducted to investigate the influence of surface roughness on the aerodynamic performance of a 1.5-stage highly loaded axial compressor at low Reynolds number. It was especially considered how the roughness Reynolds number ( k + ) affected the change of the inlet and outlet conditions, the growth of the separation bubble (LSB), the status of the limiting streamline, and the patterns of the wake. Regarding the roughness settings, five roughness magnitudes and four roughness locations were mainly studied. The results showed that at low Re, surface roughness mainly improved the stage performance by reducing the length and width of the LSB as well as the rotor tip vorticity, delaying the occurrence of three-dimensional flow separation and increasing the turbulence level near the wall. However, it also aggravated the incoordination between the subsequent stages to a certain extent, which limited further improvement of the overall aerodynamic performance. Generally, with k + increasing, the compressor aerodynamic performance improved and achieved the best at k + = 137.8 . The maximum increases in the total pressure ratio, peak efficiency, and chocked mass flow were approximately 4.01%, 5.34%, and 2.24%, respectively. In addition, for all the four roughness locations, the roughness covering from the leading edge to 50% of the axial chord length on the suction surface had a relatively evident advantage in improving the compressor peak efficiency because of the better control of the LSB and wall shear stress.

Detached Eddy Simulation of Transition in Turbomachinery: Linear Compressor Cascade

AbstractThe adaptive, ℓ2–ω delayed detached eddy simulation model was selected to simulate the flow in the V103 linear compressor cascade. The Reynolds number based on axial chord length is 138, 500. Various inflow turbulent intensities from 0% to 10% were tested to evaluate the performance of the adaptive model. The adaptive model is capable of capturing the laminar boundary layer and the large-scale perturbations inside it. The instability of large-scale disturbances signals the switch to a hybrid simulation of the turbulent boundary layer—thus, the transition front is predicted. In the case of separation-induced transition, the adaptive model, which uses eddy simulation in separated flow, can predict the separation bubble size accurately. Generally, the adaptive ℓ2–ω model can simulate the transitional separated flow in a linear compressor cascade, with a correct response to varying turbulent intensities.

Investigation on aerothermal performance of a rib-slot scheme on the multi-cavity tip of a gas turbine blade

A rib-slot blade tip is proposed to improve the film cooling performance of an unshrouded gas turbine blade tip in the current paper. The rib-slot blade tip is formed by arranging the slot into ribs of the multi-cavity squealer tip. The film cooling performance of the rib-slot blade tip is experimentally studied by using the Pressure Sensitive Paint (PSP) technique in transonic flows. The inlet Reynolds number is 370,000 based on the blade axial chord length. The exit Mach numbers is 1.05. Effects of density ratios and mass flow ratios on the film cooling characteristics are investigated, and the experiments are carried out at four mass flow ratios and three density ratios. The numerical simulation is used to predict the flow field and analyze the film cooling characteristics. The multi-cavity squealer tip consists of cavity 1-4. The results indicate that at the high mass flow ratio, the film can cover the whole squealer tip bottom surface downstream of rib-slots at three density ratios. Increasing the mass flow ratio has a slight effect on the film cooling effectiveness (η) near the tip pressure side upstream of the cavity 3. There is a low η region near the suction side of the cavity 2 and cavity 3 at low mass flow ratios, particularly for the high-density jet due to the leakage flow reattachment. A low η region is formed in the middle region of the cavity 4 at the low mass flow ratio. The blade tip with the jet shrinks the high aerodynamic loss coefficient region formed by the counter vortex, whereas, it expands the high aerodynamic loss coefficient region formed by the leakage vortex in the blade span-wise direction. The pitch-wise averaged aerodynamic loss coefficient of the blade tip with the jet is lower than that of the no film case near the shroud.