Inlet Total Temperature Research Articles

AN AERODYNAMIC INVESTIGATION OF A HIGH-PRESSURE TURBINE USING ROTOR CASING STATIC PRESSURE MEASUREMENTS AT ENGINE REPRESENTATIVE CONDITIONS WITH DIFFERENT TIP DESIGNS, TIP GAPS AND INLET TEMPERATURE PROFILES

Abstract A rotating High-Pressure (HP) turbine, in modern aircraft engines, is subjected to high cyclic thermal and mechanical loads. Shroudless turbines experience high heat loads on the rotor tip and casing, which makes them one of the life-limiting components for an engine. If the rotor tip starts to erode, then the tip clearance increases which increases the over-tip leakage flow, resulting in reduced stage efficiency and ultimately, the engine lifetime. The Oxford Turbine Research Facility (OTRF) is a high- speed rotating transient test facility that allows unsteady aerodynamics and heat transfer measurements, at engine representative conditions. This paper presents an experimental investigation, of the HP turbine rotor casing aerodynamics, performed in the OTRF. Steady and unsteady casing static pressure measurements were acquired, using pneumatic lines and high-frequency response (∼ 100 kHz) Kulite pressure transducers respectively. The single-stage HP turbine consisted of cooled vanes (featuring film and trailing edge slot cooling), and uncooled rotor blades. Three different tip designs (two squealers and one flat) were tested, at two tip gaps, and with two inlet temperature profiles that included a spatially uniform total temperature profile and an engine representative HP turbine inlet total temperature profile. In parallel, a numerical investigation of the experimental test cases is presented using unsteady Computational Fluid Dynamics (CFD) predictions. The experimental results from this study provide a unique dataset to validate numerical models.

Aerodynamic and aeroacoustic sensitivities of contra-rotating open rotors to operational parameters

Abstract Open rotors can play a critical role towards transitioning to a more sustainable aviation by providing a fuel-efficient alternative. This paper considers the sensitivity of an open-rotor engine to variations of three operational parameters during take-off, focusing on both aerodynamics and aeroacoustics. Via a sensitivity analysis, insights to the complex interactions of aerodynamics and aeroacoustics can be gained. For both the aerodynamics and aeroacoustics of the engine, numerical methods have been implemented. Namely, the flowfield has been solved using unsteady Reynolds Averaged Navier Stokes and the acoustic footprint of the engine has been quantified through the Ffowcs Williams-Hawking equations. The analysis has concluded that the aerodynamic performance of the open rotor can decisively be impacted by small variations of the operational parameters. Specifically, blade loading increased by 9.8% for a 5% decrease in inlet total temperature with the uncertainty being amplified through the engine. In comparison, the aeroacoustic footprint of the engine had more moderate variations, with the overall sound pressure level increasing by up to 2.4dB for a microphone lying on the engine axis and aft of the inlet. The results signify that there is considerable sensitivity in the model and shall be systematically examined during the design or optimisation process.

Uncertainty quantification of the inlet boundary conditions in a supercritical CO2 centrifugal compressor based on the non-intrusive polynomial chaos

Given the sensitivity of supercritical carbon dioxide centrifugal compressors performance to inlet boundary parameter fluctuations, this study focuses on the uncertainty quantification of the inlet boundary conditions. Non-intrusive polynomial chaos surrogate models were developed for both performance parameters and flow field of a supercritical carbon dioxide centrifugal compressor. Sensitivity and correlation analyses revealed that inlet total temperature is the primary influencing parameter. Fluctuations of 1.37 % in total temperature and 6.75 % in total pressure result in changes of 40.94 %, 18.57 %, and 1.77 % in mass flow rate, total pressure ratio, and total to total isentropic efficiency, respectively. Flow field statistical analysis shows that inlet uncertainties nonlinearly impact the dryness fraction and relative Mach number distributions near the blade leading edge, while the density in the condensation zone on the suction side of the leading edge is less affected. Error bands in static pressure on blade surfaces highlight significant variations, particularly at 90 % spanwise on the blade suction side, where shocks and supersonic flow induce prominent disturbances. This study offers a preliminary methodology and theoretical guidance for robust aerodynamic optimization and operational strategies of supercritical carbon dioxide centrifugal compressors.

Degradation performance rapid prediction and multi-objective operation optimization of gas turbine blades

The performance degradation of high-temperature blades will affect the safety and efficiency of operation. In response to issues such as ignoring strength characteristics, slow degradation performance prediction, insufficient applicability and accuracy of traditional proxy models, and long operation optimization cycle, a degradation performance prediction model of turbine blades based on the back propagation neural network is established in this research. Furthermore, the efficiency-stress multi-objective operation optimization method based on genetic algorithm is proposed. The results show that the R-square value of the degradation performance neural network prediction model is greater than 0.9 for predicting characteristic parameters. The prediction speed is improved by 104 orders of magnitude compared to the traditional computational fluid dynamics method. The influence mechanisms of turbine outlet flow rate, cooling gas inlet total temperature, and cooling gas flow rate on the degradation performance are revealed. Under fouling extreme conditions and corrosion/wear extreme conditions, the operating parameters of turbine blades are optimized. The maximum values of the efficiency improvement, power increase and maximum von Mises stress reduction are 0.74 %, 2.13 MW and 43 MPa, respectively. This study can provide an effective and novel method for real-time degradation performance prediction and rapid operating condition optimization of turbine blades.

Unsupervised correction algorithm of engine inlet total temperature

Unsupervised correction algorithm of engine inlet total temperature

Experimental and numerical investigations on controlled parameter selection methods for kerosene-fueled scramjet

Abstract In order to determine the optimal wall pressure location as the controlled parameter for kerosene-fueled scramjet, a combined approach of numerical simulation and experimental analysis was employed. The results indicate that with an increase in the equivalence ratio of kerosene, the combustor exhibits consecutive modes, including pure scramjet mode, dual-mode scramjet mode, dual-mode ramjet mode, and pressure disturbance to the isolator inlet mode. The peak pressure ratios corresponding to the mode transition boundaries are 1.95, 3.26, and 3.57, with respective isolator outlet Mach numbers of 1.62, 1.0, and 0.74. The equivalence ratio corresponding to mode transition increases with an increase in the combustor inlet total temperature. Considering the dynamic characteristics of wall pressures at different locations, the position at x/L = 0.46 demonstrates the best linearity, highest sensitivity, and greater stability. Hence, it is the most suitable choice as the controlled parameter for fuel flow in this kerosene-fueled scramjet.

Stall and stability enhancement mechanisms of transonic compressor with inlet total temperature distortion

Inlet total temperature distortion refers to the nonuniform distribution of the total temperature at the inlet of the aero-engine, which is one of the external destabilizing factors that have the most remarkable impact on the stability of an aero-engine. In this study, the stalling process of the Darmstadt transonic compressor is investigated under the total temperature distortion of 180 deg circumferential extent and 500 K intensity by full-annulus unsteady numerical simulation. The analysis shows the addition of inlet total temperature distortion deteriorates the compressor stall margin from 22.7% to 17.0% with a considerable decrease in the pressure ratio. The type of compressor stall inception under temperature distortion conditions remains spike. Pressure perturbations and radial vortex formation are first clearly detected when the rotor rotates into the high temperature distortion region. The circumferential propagation of the stall cells under the total temperature distortion is 66% of the rotor speed, which is faster than that under uniform condition whose value is 44% of the rotor speed. The optimized casing treatment (CT) has extended the stall margin of the rotor without producing efficiency loss under uniform condition. The anti-distortion ability of CT is first verified despite its negative impact on efficiency under total temperature distortion. The adoption of CT could obviously push the shock wave into the blade passage under total temperature distortion condition. In addition, it can reduce the tip blade loading, thus removing the low Mach number area in the tip region, while increasing the blade loading below 80% span.

Investigation on high-dimensional uncertainty quantification and reliability analysis of aero-engine

Accurate quantification of uncertainties is paramount in the contemporary design of aero-engines, demanding substantial computational resources. This investigation introduces an efficient parallel framework for uncertainty quantification, significantly curtailing sample requirements while upholding computational precision. Furthermore, this research tackles the constraints intrinsic to traditional reliability analyses, which often disregard uncertainty, by conceiving an algorithm tailored to aero-engine reliability analysis and lifespan prediction that deliberately accounts for uncertainties. The findings from the uncertainty quantification disclose that the most pivotal factor contributing to the variability in leakage flow rate is unequivocally the variance index of the tip clearance deviation, demonstrating a substantial variance index of 81.52 %. Concerning cooling performance, a remarkable shift is evident, with inlet total temperature fluctuation eclipsing the tip clearance deviation in terms of significance. This transformation is accentuated by a variance index of 54.52 %. Furthermore, the film cooling effectiveness displays notable variability within the uncertainty framework, underscoring that the blades may be subjected to more severe thermal corrosion and thermal fatigue during operational conditions than initially envisioned. Reliability calculations elucidate that the probability of blade failure experiences the steepest ascent after approximately 15,500 h of operation, necessitating mandatory overhaul at this juncture.

Loss reduction in the compressor corner region via blade cooling

With Reynolds number (Re) decreasing to a critical value (approximately 2.0 × 105), the laminar-to-turbulent transition on the blade/endwall surface accelerates the loss generation in the corner region of compressors. Thus, how to effectively control the three-dimensional turbulent mixing loss at a low Re is of great importance. In this study, the mechanism of loss variation in a high-speed compressor cascade as Re decreases from 5.6 × 105 to 1.5 × 105 is investigated by large eddy simulations (LESs). Then, an isothermal cooled wall condition with a constant temperature of 0.8T01 (T01 is the inlet total temperature) is specified on the blade surface to explore the potential of loss reduction via wall cooling at Re=1.5 × 105. At Re=5.6 × 105, the boundary layers on the blade suction surface strongly interact with the endwall boundary layers and multiple passage vortex structures near the spanwise location of x/H = 0.3, inducing a large-scale concentrated shedding vortex (CSV) and determining the loss generation rate in the corner region. With Re decreasing to 1.5 × 105, the blade boundary layers grow much thicker due to the delayed transition. Detailed loss classification result indicates that the stronger interaction between the thickened blade boundary layers and secondary flow accelerates the local generation of turbulent fluctuations, which is mainly responsible for the performance deterioration in the corner region at a low Re. As the blade surface is cooled, the growth of blade boundary layers and the migration of secondary flow in the blade passage can be controlled simultaneously. Thus, the interaction between the blade boundary layers and secondary flow near the spanwise location of x/H = 0.3 is weakened, leading to a reduction in the overall total pressure loss by 13.3%. The results hope to guide in controlling the loss in the corner region of compressors operating at a low Re.

Comprehensive Aerothermal Investigation of Turbine Blade Multicavity Squealer Tip Using a Novel Methodology With Uncertainty Quantification

Abstract Ascertaining the uncertainty in the aerothermal performance of blade tips is crucial, as it represents the most delicate component of modern gas turbines. In this research article, a novel and efficient approach is proposed for quantifying uncertainties in aerothermal performance using a combination of universal kriging, polynomial chaos expansions, and Smolyak sparse grid technology. This method was applied to investigate the aerothermal performance of a high-pressure gas turbine rotor blade tip with high-dimensional robustness. The outcomes of the uncertainty quantification calculation reveal that the downstream total pressure loss coefficient and leakage flowrate increase under normal-speed (subsonic) and high-speed (transonic) conditions. The key uncertainty input that affects the aerodynamic performance of normal-speed and high-speed squealer tip is inlet total pressure fluctuation, with a variance index on the leakage flowrate of normal-speed and high-speed squealer tip of up to 73.92 and 83.85%, respectively. The study suggests that it is more important to control the operating conditions fluctuation than the cavity depth machining accuracy for aerodynamic performance robustness, which applies to both normal-speed and high-speed squealer tips. In line with the aerodynamic performance, the heat flux of normal-speed and high-speed squealer tip increases during operation. Notably, the sensitivity of high-speed squealer tip aerodynamic performance to operating condition fluctuations increases compared to the normal-speed squealer tip, necessitating active intervention for fluctuations in operating conditions at a higher cost for the high-speed squealer tip. The sensitivity analysis results indicate that the inlet total temperature fluctuation is the key parameter that controls the normal-speed and high-speed squealer tip heat flux uncertainty. Finally, it is worth noting that while the use of ribs can effectively enhance the robustness of blade tip heat transfer performance, the heat flux near the root of the ribs fluctuates significantly, which may further increase the thermal fatigue tendency in this region during actual operation.

A novel combined model for energy consumption performance prediction in the secondary air system of gas turbine engines based on flow resistance network

Predicting the comprehensive energy consumption is challenging for the complicated secondary air system. To solve this problem, a universal flow resistance element has been constructed to describe different components more accurately. And a novel flow resistance network is proposed to calculate the system's energy consumption. This network can be analyzed by the flow resistance characteristics of its components to estimate the energy consumption of each component and the energy loss of the system. Furthermore, a flow resistance parameter is defined as a unique factor to determine the magnitude of flow resistance for each component. Compared to experimental and numerical results, both models demonstrate sufficient accuracy in calculating mass flow rates and entropy increments, with the maximum deviation less than 1.2%. The above two models are also applied to predict flow losses of elements with changes in pressure and inlet total temperature, with the maximum deviation less than 4.5%. Based on the flow resistance network, the aerodynamic performance of each element can be easily computed by the inlet and outlet boundary of this system, regardless of whether the elements are arranged in series or parallel.

Performance estimation of multi-stage cooled axial flow turbines under choked conditions

Development of a fast and accurate tool to estimate the performance of axial flow turbines in all range of working conditions including choked region is of great interest in the design steps of turbines. Mean-line (one-dimensional) analysis is one of the solutions that is computationally cost-effective along with producing accurate results. However, the development of a mean-line model that able to solve the flow in a choked turbine (one-row chocking and multi-row choking) is a challenging problem. This study presents a new algorithm for mean-line analysis of axial flow turbines with a set of well-posed boundary conditions, i.e., the inlet total pressure, the inlet total temperature, inlet flow angle, and outlet static pressure and calculates the flow condition through the turbine row-by-row. The proposed algorithm is implemented in an object-oriented C++ code named Axial Turbine Simulator (ATSim). The results of ATSim program is validated and verified under various working conditions, e.g., low and high-pressure ratios (choking condition), for four different turbine configurations including two uncooled single-stage turbines, an uncooled two-stage turbine, and a cooled three-stage turbine. ATSim provide good consistency with previous experimental and numerical studies. Moreover, the results show that ATSim is capable of estimating the performance of the turbine in a very short time over a wide range of working conditions, including choking (one-row choking or multi-row choking) region.

Deep reinforcement learning method for turbofan engine acceleration optimization problem within full flight envelope

In order to solve the multidimensional restricted optimization problem of aeroengine acceleration process and improve the acceleration performance of full envelope, a design method of aeroengine full envelope acceleration controller based on deep reinforcement learning is proposed. Firstly, the thermo-aerodynamic principle of the engine acceleration process is analyzed, and the mathematical analytical model of the multidimensional restricted optimization problem of the turbofan engine acceleration process is established. On this basis, the design method of acceleration controller based on twin delayed deep deterministic policy gradient is studied. The proportional integral controller is designed to pursue the asymptotic stability at the acceleration destination and stable control is achieved through switching. Additionally, the flight envelope is divided into regions by clustering to reduce the span of data changes, so as to enhance the network convergence ability of deep reinforcement learning. Finally, based on the similarity conversion method of equal engine inlet total temperature, the acceleration controller of full flight envelope is realized. The digital simulation results demonstrate that compared with the traditional proportional integral differential controller, the acceleration controller based on twin delayed deep deterministic policy gradient shortens the acceleration time up to 48.33%, and the control effect in a large flight envelope (H=0∼10 km, Ma=0∼1.6) has been verified.

Performance assessment for a novel supersonic turbine engine with variable geometry and fuel precooled: From feasibility, exergy, thermoeconomic perspectives

Aiming at the problem of turbine engine insufficient thrust caused by low efficiency at high Mach number, a three-axis variable geometry dual-mode turbine engine scheme based on ammonia and n-decane equivalent dual fuel precooling and adjustable low-pressure (LP) turbine guide vane is proposed, and parameters such as specific thrust price, relative thrust economic ratio, volumetric specific fuel consumption are defined to scientifically evaluate the feasibility of the scheme. The ammonia and n-decane dual fuel method is used to expand the efficient working range of the precooler. On the one hand, in the low speed section, the engine operates in the turbofan mode, adjusting the LP turbine guide vane can offset the adverse impact of the precooler low efficiency on the engine thrust performance under the low inlet total temperature, and improve the engine economy. On the other hand, the high speed section operates in the turbojet mode, and the equivalent fuel precooling combined with the LP turbine guide vane adjustment is conducive to improving the working efficiency and thrust performance of the turbojet engine. Feasibility, exergy and thermoeconomic analysis indicate that under different operating modes, due to different control plans, the influence of fuel ratio and heat transfer efficiency changes of dual fuel precooler on engine performance is opposite. For the typical climbing process with H = 0, Ma = 0 to H = 30 km, Ma = 5, this scheme can significantly increase the engine thrust, with a maximum increase of about 41%, but it increases the volume fuel consumption, reduces the engine exergy efficiency, exergy sustainability, and thermoeconomic under the turbofan mode. Therefore, the variable geometry turbine engine scheme based on dual fuel precooling and adjustable LP turbine guide vane has high application potential in supersonic commercial flight under the premise of slightly losing economy.

Investigation of rotating detonation gas turbine cycle under design and off-design conditions

As cycle efficiency of gas turbine can be significantly improved due to the pressure gain characteristic of rotating detonation combustion, a cycle calculation model of rotating detonation gas turbine based on component characteristic curves of a single shaft gas turbine of 25 MW and limited maximum flow capacity of rotating detonation combustor was established to investigate the variations of cycle performance and component characteristics under design and off-design conditions. Thereafter, sensitivity analysis of component parameters was conducted to explore the comprehensive reason for change of cycle efficiency increment. The results demonstrated that compared with those of traditional gas turbine, higher cycle performance and lower compressor pressure ratio were gained in rotating detonation gas turbine. Increment of cycle efficiency and cycle net power increment significantly increased with the drop of load, which reached to 0.0248 and 3.261 MW when turbine inlet total temperature was 1450 K and increased to 0.0539 and 4.008 MW when turbine inlet total temperature was 1279 K. Cycle performance was significantly affected by the changes of compressor pressure ratio, compressor efficiency, equivalence ratio of RDC and turbine efficiency, among which turbine efficiency rather than compressor outlet total temperature had the most significant effect on cycle efficiency increment under the calculated conditions.

Impact investigation of inlet environmental changes on the rated condition performance of a high-pressure compressor

Abstract The variations of inlet environment parameters can make significant effects on the compressor performance. This paper aims to investigate the effects of inlet total pressure and total temperature changes on the rated condition performance of a nine-stage HPC. Different cases of total pressure and total temperature boundary conditions at this compressor inlet are studied by 3-D numerical simulations with experimental validations. The numerical results confirm that the variations of inlet total pressure and total temperature make different effects on the rated condition performance of compressor. The overall performance parameters, such as the corrected mass flow and isentropic efficiency, will increase with inlet total pressure increasing and decrease with inlet total temperature increasing by different change rules. The flow similarity is also investigated by comparing the calculated results of critical quantities in different cases. The results indicate that the rising inlet total pressure can increase the Reynolds number and it is beneficial to reduce the viscous influence so that it is available to improve the performance; the rising inlet total temperature can decrease both the specific heat ratio and Reynolds number so that it will lead to the compressor performance decline inevitably.

Robustness Analysis on Aerothermal Performance of the Winglet Squealer Tip—Part 1: Aerodynamic Performance

Abstract Accurately assessing the robustness of the aerothermal performance of the blade tip is important considering that uncertainty is inevitable in the actual operation of turbines. However, the conventional uncertainty quantification methods are computationally inefficient for such an expensive black-box problem as turbine aerothermal performance prediction. In this paper, an efficient framework that is based on the combination of the sparse polynomial chaos expansion (PCE) and universal Kriging (UK) metamodel is applied to the uncertainty quantification of the effect of the conventional squealer tip and three different winglet squealer tips on the aerodynamic performance of the GE-E3 rotor blade tip. However, the inlet total pressure, inlet total temperature, and inlet flow angle are considered to flow condition uncertainty parameters and tip clearance is considered a geometrical uncertainty parameter. According to the results of the uncertainty quantification, in actual operation, although the setup of the winglet structure can still reduce the leakage flowrate, its effect will be much lower than predicted by deterministic calculations. The parameter that has the greatest influence on the uncertainty of the aerodynamic performance of the four tip structures is the tip clearance. Therefore, the geometric accuracy of the tip clearance should be strictly ensured in the turbine blade assembly and marching process. The uncertainty quantification calculations reveal that there is an antagonistic relationship between the pressure side cavity and suction side cavity on the aerodynamic performance uncertainty of the blade tip, which indicates a reasonable ratio of pressure side cavity and suction side cavity can make the fluctuation of the aerodynamic performance of the pressure side cavity vortex and suction side cavity vortex completely cancel, and thus design the winglet squealer tip with strong aerodynamic performance robustness.

Robustness Analysis on Aerothermal Performance of the Winglet Squealer Tip—Part II: Heat Transfer Performance

Abstract This article is the second part of a robustness analysis of the aerothermal performance of the winglet structure. A challenging analytical test function and an engineering test are considered to further investigate the response performance of the efficient uncertainty quantification framework proposed in Part 1. Then, a series of original visualizing uncertainty quantities are proposed in this article and applied to the study of uncertainty quantification of the winglet structure. Finally, this efficient framework is applied to the uncertainty quantification of the effect of the conventional squealer tip and three different winglet squealer tips on the heat transfer performance of the GE-E3 rotor blade. According to the results of the uncertainty quantification calculation, in actual operation, the setting of the winglet structures will diminish rather than increase the heat transfer performance of the blade tip. This conclusion is completely opposite to the prediction of the deterministic calculation. The heat flux increase and standard deviation of squealer tip with pressure-side winglet are the highest among the four tip structures, which means that the robustness of heat transfer performance of squealer tip with pressure-side winglet is the worst. The parameter that has the greatest influence on the uncertainty of the heat transfer performance of the four tip structures is the tip clearance. But the influence of the inlet total temperature fluctuation must also be taken into account. So a satisfactory control system should be designed for the actual operation of the gas turbine so that the fluctuation of inlet total temperature can be attenuated rapidly. A positive correlation between the heat flux of the blade tip mean value and the standard deviation is revealed by the uncertainty quantification, which implies that reducing the heat flux of the blade tip mean value in the robust design of the blade tip tends to reduce the heat flux fluctuation as well. Therefore, the objective of the robust design of the blade tip can be either one of reducing the mean value of heat flux of the blade tip mean value or reducing the heat flux of the blade tip standard deviation without multi-objective optimization. It is worth noting that, like the aerodynamic performance uncertainty, there is an antagonistic relationship between the pressure-side cavity and suction-side cavity on the heat transfer performance uncertainty of the blade tip. Therefore, a reasonable ratio of pressure-side cavity to suction-side cavity in the turbine design can also lead to a blade tip with strong heat transfer performance robustness.

Full-envelope acceleration control method of turbofan engine based on variable geometry compound adjustment

In order to reduce the influence of compressor stability margin constraint and enhance the full envelope acceleration performance of turbofan engine, a full envelope acceleration control method of turbofan engine based on variable compressor intermediate stage bleed and adjustable low-pressure (LP) turbine guide vane is proposed. Firstly, a high confidence turbofan engine component level model is established, which is capable of modifying the compressor adiabatic efficiency caused by variable intermediate stage bleed and adjusting the LP turbine guide vane adaptively. Then, the main combustion chamber fuel flow, compressor intermediate stage bleed flow and LP turbine guide vane angle are selected as the optimization variables. Meanwhile, the PSO-SQP (Particle Swarm Optimization-Sequential Quadratic Programming) method is adopted to obtain the acceleration compound control schedule. Finally, the full envelope compound adjustment acceleration control plan is available based on the similarity conversion theory of consistent engine inlet total temperature. The simulation results demonstrate that the acceleration control method based on variable geometry compound adjustment can significantly promote engine acceleration performance in the left flight envelope, shorten the maximum acceleration time by more than 50%, and dramatically improve the stability margin of the compressor during acceleration phase. In the right flight envelope, the variation law of the minimum surge margin during acceleration procedure almost coincides with the engine inlet isotherm. It proves that the method proposed possesses good applicability and feasibility in the full envelope.

Robustness Analysis on the Aerothermal Performance of Turbine Blade Squealer Tip

Abstract An improved efficient uncertainty quantification (UQ) analysis framework is proposed by the combination of sparse polynomial chaos expansion (PCE) and universal Kriging (UK) metamodel to obtain the surrogate model (UK-PCE). Moreover, a challenging analytical test function and an engineering test are considered to investigate the response performance of UK-PCE method. The results show that the UK-PCE method reduces the computational cost by more than 70% in comparison to the typical PCE method. Then this method was applied to the UQ of the aerodynamic and heat transfer performance of GE-E3 rotor blade squealer tip. Additionally, a series of uncertainty quantities visualization methods based on the data mining method, parallel computing method, and Delaunay triangulation method is proposed to reveal more enlightening uncertainty phenomena in the actual operation. The results of UQ show that under the influence of uncertain inputs, the leakage flow rate and downstream entropy increase will be significantly increased. The statistical average of tip heat flux has increased by 8.56% relative to the design value, and the probability of it deviating from the design value by 10% is as high as 43.27%. In addition, the three-dimensional tip heat flux deviation distributions calculated by the proposed uncertainty quantities visualization method reveal a coupling of the hot corrosion and thermal fatigue of the squealer tip. It is also indicated that under the influence of the uncertain inputs, there is a marked increase in blade tip flux, and the blade tip flux deviation has been maintained at a high value, about 13.0%. The results of sensitivity analysis show that the largest contributor to the uncertainty of the blade tip aerodynamic performance is the tip clearance deviation and its variance index to the uncertainty of leakage flow rate and downstream entropy increase is as high as 88.21% and 62.63%. Therefore, the geometric accuracy of the tip clearance should be strictly ensured in the turbine blade assembly and marching process. The influence of the inlet total temperature deviation on the uncertainty of the heat transfer performance of the squealer tip must also be taken into account. So a satisfactory control system should be designed in the actual operation of the gas turbine to make sure that the fluctuation of inlet total temperature can be attenuated rapidly.