High Pressure Turbine Rotor Blade Research Articles

High pressure turbine rotor blades heating duration experimental estimates prior to the bypass turbofan engine start for reducing thermal stresses

Reducing the durability cost of an aircraft product is an issue either addressed at the design stage or causing significant design modifications. Turbine preheating of a gas turbine engine (GTE) allows for the thermal stress of the rotor blades (RB) to be reduced at the engine start without making design changes, but only by implementing the engine heating technology into the operational process. Values of thermal stresses on rotor blades of a high-pressure turbine of a bypass turbofan engine (TFE) with and without heating allow us to determine the change in the total HPT RB damage rate. In the concept of preheating a GTE prior to the start in order to comply with the preheating technology, it is necessary to know the duration within which the RB will heat up to the required temperature. Thus, the research objective, presented in the paper, is to empirically determine the HPT RB heating time, using thermocouples and pyrometers on a full-scale body depending on the methods of air supply for heating and rotor spinning. A distinctive feature of this work is the application of the empirical approach to determine HPT RB heating time to evaluate the feasibility of the GTE preheating technology application prior to the start and the selection of the most efficient heating method according to the duration criterion. Several methods of engine heating prior to the start, using different sets of equipment and the method of supplying hot air to the turbine, were considered. The results of RB heating time measurements made it possible to establish the method of heating with minimal time expenditure prior to the engine start.

Heat Transfer Analysis of Damaged Shrouded High-Pressure Turbine Rotor Blades

Due to the increasingly high turbine inlet temperatures, heat transfer analysis is now, more than ever, a vital part of the design and optimization of high-pressure turbine rotor blades of a modern jet engine. The present study aimed to find out how shape deviation and in-service deterioration affect heat exchange patterns on the rotor blade. The rotor geometries used for this analysis are represented by a set of high-resolution 3D structured light scans of blades with the same number of in-service hours. An automatic meshing technique was employed to generate high-resolution meshes directly on the scanned rotor geometries, which captured all the surface features with high fidelity. Steady-state 3D RANS flow simulations with a k-ω SST turbulence model were conducted on a one-and-a-half stage computational domain of the scanned geometries. First, the distribution of the heat transfer coefficient was calculated for each blade; then, a correlation was sought between the heat transfer coefficient and parametrized shape deviation, to assess the impact of each parameter on HTC levels.

High-Fidelity Computational Fluid Dynamics Analysis of In-Serviced Shrouded High-Pressure Turbine Rotor Blades

Abstract In-service deterioration can lead to undesired shape variations on high-pressure turbine rotor blades. This can have a significant impact on efficiency, power generation, and component life. Loss of power production from the high-pressure turbine rotor causes engine over-throttling to compensate for the lower performance. This will in turn worsen the operating conditions and ultimately reduce the life of the component. The aim of this study is to provide a high-fidelity flow simulation of in-serviced shrouded high-pressure turbine (HPT) blades of a modern Jet Engine. Shape variation effects on the aerodynamic performance of several shrouded HPT blades with a different number of in-service hours have been investigated. In order to establish a digital model of the shape variation, a novel reverse-engineering procedure is carried out to come up with a parametrized definition of each blade's variations from nominal including any observable damage. The investigation is conducted by means of an in-house, full 3D steady-state Reynolds-averaged Navier–Stokes (RANS) simulation of the flow around a series of damaged rotor blade geometries, which are obtained through high-resolution optical blue-light “Gesellschaft für Optische Messtechnik” (GOM) scans. The analysis shows that the aerodynamic performance of the HPT rotor blades under investigation is primarily sensitive to shroud damage, which is found to account for efficiency losses often greater than 3%, and for more than 80% of the total performance loss. A secondary role on efficiency is found to be played by the blade shape deviation. A highly linear correlation is found between HPT stage efficiency and a combination of shroud damage parameters.

Theoretical and experimental evaluations of pre-swirl rotor-stator system with inner seal bypass configuration for turbine performance improvement

Pre-swirl system has strong potential to improve the cooling performance and service life of high-pressure turbine rotor blades in gas turbine engines. This study is to evaluate and improve the temperature drop and power consumption characteristics of the pre-swirl system with multi-in and multi-out seal flow by theoretical and experimental analysis. Then, a high-speed rotating test rig of the pre-swirl rotor-stator system is established, considering the inner seal by-pass configuration to prevent the flow of hot gas into the pre-swirl cavity. In addition, the rotor-stator domain parameters in the pre-swirl system test rig are accurately measured at high-speed operating conditions. Three experimental schemes of pre-swirl systems are proposed for comparison under different operating conditions. The experimental results show that the temperature drop of the prototype pre-swirl system is in the range of 9.28–20.78 K, with the power consumption of the system ranging between −0.23 and −6.67 kW and the system pressure ratios being around 1.3–1.9. Because of the inner seal inflow ratio ranging within 0–15%, the system temperature drops for the improved pre-swirl system built-in inner seal by-pass passages notably reach a value between 12.12 K and 25.54 K with an average system power consumption within −0.41 to −8.21 kW. Thus, the improved pre-swirl system arrangement has dual advantages of a higher cooling effect and lower power consumption.

Exploring Topology Optimization for High Pressure Turbine Blade Tips

Abstract This work presents the aerodynamic topology optimization of high-pressure turbine rotor blade tips. Before carrying out the topology optimization on the blade tip, some initial tip design studies were carried out. A winglet tip was optimized first and it was found that the optimum winglet design features a combination of small and largest overhangs possible that increase the aerodynamic efficiency by 1.40% compared to the datum design. Second, a radial basis function-based parametrization was set up to allow the creation of a single squealer rim on the datum blade’s tip that could move position in the circumferential direction. The optimum case proved to increase efficiency by 0.46% compared to the flat datum tip of the same tip gap. After that, a combination of winglet and topology free squealer tips was investigated via topology optimization. The winglet tip was created as in the winglet-only optimization cases and topology free squealer walls were parametrized and created using mapping of a radial basis function surface. It was shown that the radial basis function surface-based parametrization creates a very flexible design space containing novel squealer topologies. Combining both winglet and novel squealer topology optimization, better designs than the flat tip winglet can be achieved. However, because of the flexibility of the design space, gradient-based methods were found to struggle to reach an optimum solution. This was resolved by optimizing the most promising design subspace.

The required aerodynamic simulation fidelity to usefully support a gas turbine Digital Twin for manufacturing

With the imminent digitalisation of the manufacturing processes of gas turbine components, a large volume of geometric data of as-manufactured parts is being generated. This geometric data can be used in aerodynamic simulations to predict component performance. Both the cost and accuracy of these simulations increase with their fidelity. To efficiently exploit Digital Twin technology, one must therefore understand how realistic the aerodynamic simulations need to be to give useful performance predictions. This paper considers this issue for a sample of scrapped high-pressure turbine rotor blades from a civil aero engine. The measured geometric data was used to build aerodynamic models of varying degrees of realism, ranging from quasi-three-dimensional blade sections for an Euler solver to three-dimensional, multi-passage and multi-stage Reynolds-Averaged-Navier-Stokes models. The flow near the tip of these shrouded blades is sensitive to manufacturing variability and can switch between two quasi-stable horseshoe vortex modes. In general, capacity and exit flow angle can be adequately predicted by three-dimensional, single-passage calculations: averaging single-passage calculations gives a good prediction of the multi-passage behaviour. For efficiency and stage loading, the approach of averaging single-passage calculations is less accurate as the multi-passage behaviour requires an accurate prediction of the horseshoe vortex modes.

Film Cooling Characteristics of the Suction Surface of a High Pressure Turbine Rotor Blade

Film Cooling Characteristics of the Suction Surface of a High Pressure Turbine Rotor Blade

Physics-Based Part Orientation and Sentencing: A Solution to Manufacturing Variability

Abstract Casting deviations introduce geometric variability that impacts the aerodynamic performance of turbomachinery. These effects are studied for a high-pressure turbine rotor blade from a modern aero-engine. A sample of 197 blades were measured using structured-light three-dimensional scanning, and the performance of each blade is quantified using Reynolds-averaged Navier–Stokes (RANS) simulations. Casting variation is typically managed by applying geometric tolerances to determine the suitability of a component for service. The analysis demonstrates that this approach may not be optimal since it does not necessarily align with performance, in particular the capacity and efficiency. Alternatively, functional acceptance based on the predicted performance of each blade removes the uncertainty associated with geometric tolerancing and gives better performance control. Building on these findings, the paper proposes a method to set the orientation of the fir-tree, which is machined after casting. By customizing the alignment of each blade, performance variability and scrap rates can be significantly reduced. The method uses predictions of performance to reorient the castings to compensate for manufacturing-induced errors, without changing the design-intent blade geometry and with minimal changes to the manufacturing facility.

Gas temperature measurement of high-pressure turbine rotor blades during bench tests of a core engine

Gas temperature measurement of high-pressure turbine rotor blades during bench tests of a core engine

Effects of rising angle on upstream blades and intermediate turbine duct

With the improvement of requirement, design and manufacture technology, aero-engines for the future are characterized by further reduction in fuel consumption, cost, but increment in propulsion efficiency, which leads to ultra-high bypass ratio. The intermediate turbine duct (ITD), which connects the high pressure turbine (HPT) with the low pressure turbine (LPT), has a critical impact on the overall performances of such future engines. Therefore, it becomes more and more urgent to master the design technique of aggressive, even super-aggressive ITDs. Over the last years, a lot of research works about the flow mechanism in the diffuser ducts were carried out. Many achievements were reported, but further investigation should be performed. With the aid of numerical method, this paper focuses on the change of performance and flow field of ITD, as well as nearby turbines, brought by rising angle (RA). Eight ITDs with the same area ratio and length, but different RAs ranges from 8 degrees to 45 degrees, are compared. According to the investigation, flow field, especially outlet Ma of swirl blade is influenced by RA under potential effect, which is advisable for designers to modify HPT rotor blades after changing ITD. In addition to that, low velocity area moves towards upstream until the first bend as RA increases, while pressure loss distribution at S2 stream surface shows that hub boundary layer is more sensitive to RA, and casing layer keeps almost constant. On the other hand, the overall total pressure loss could keep nearly equivalent among different RA cases, which implies the importance of optimization.

Application of conjugate heat transfer and fluid network analysis to improvement design of turbine blade with integrated cooling structures

A combination of fluid network analysis method with conjugate heat transfer are applied to the improvement design of the integrated cooling structures in a high-performance turbine blade, coupled with the 3D viscous solver for the gas flow field. By comparison with the experimental results of open literatures, the methodology developed is numerically validated. For a high-pressure turbine rotor blade, it is used to rapidly predict and evaluate the aerodynamic and heat transfer performances of its integrated inner cooling structures. According to the computation results, three ways are definitely proposed for the improvement design, including the adjustment of the coolant flow mass entering into the front and rear cavities in a more appropriate flow mass ratio, the improvement of the turning geometries in serpentine channels to minimize the inner coolant flow resistance, and the adjustment of the local cooling structure dimension according to the high temperature region on outer surface of blade. Through the verification of the fully 3D conjugate heat transfer simulation for the fields of gas flow, solid blade and coolant flow, it shows that the maximum temperature on rotor blade surface is reduced obviously, the temperature distribution becomes more uniform after improvement, and the inlet parameters of cooling cavities are matched more reasonably. It is concluded that in this paper the fluid network combined with conjugate heat transfer significantly shortens the aerodynamic and heat transfer design cycle for the turbine blade with integrated cooling structures.

Condition assessment of gas turbine blades and coatings

This paper brings out the metallurgical investigation that was carried out on discoloured low pressure turbine rotor blade (LPTR) for its platinum modified aluminium coating integrity and overheated high pressure turbine rotor blade (HPTR) for its blade material and coating integrity. LPTR blades were investigated as they were for the requirement of reuse. Heat tinting, EDS and XRD were carried out to present the harmonized inference that the coating was in good condition. HPTR blades were sectioned to study under SEM. Regions of platinum modified aluminide coating degradation and DS CM 247 LC base material cracking were observed which make the particular lot unusable.

PIV Maps of Tip Leakage and Secondary Flow Fields on a Low-Speed Turbine Blade Cascade With Moving End Wall

New, detailed flow field measurements are presented for a very large low-speed cascade representative of a high-pressure turbine rotor blade with turning of 110deg and blade chord of 1.0m. Data were obtained for tip leakage and passage secondary flow at a Reynolds number of 4.0×105, based on exit velocity and blade axial chord. Tip clearance levels ranged from 0% to 1.68% of blade span (0% to 3% of blade chord). Particle image velocimetry was used to obtain flow field maps of several planes parallel to the tip surface within the tip gap, and adjacent passage flow. Vector maps were also obtained for planes normal to the tip surface in the direction of the tip leakage flow. Secondary flow was measured at planes normal to the blade exit angle at locations upstream and downstream of the trailing edge. The interaction between the tip leakage vortex and passage vortex is clearly defined, revealing the dominant effect of the tip leakage flow on the tip end-wall secondary flow. The relative motion between the casing and the blade tip was simulated using a motor-driven moving belt system. A reduction in the magnitude of the undertip flow near the end wall due to the moving wall is observed and the effect on the tip leakage vortex examined.

Numerical Investigation on the Influence of Hot Streak Temperature Ratio in a High-Pressure Stage of Vaneless Counter-Rotating Turbine

The results of recent studies have shown that combustor exit temperature distortion can cause excessive heat load of high-pressure turbine (HPT) rotor blades. The heating of HPT rotor blades can lead to thermal fatigue and degrade turbine performance. In order to explore the influence of hot streak temperature ratio on the temperature distributions of HPT airfoil surface, three-dimensional multiblade row unsteady Navier-Stokes simulations have been performed in a vaneless counter-rotating turbine (VCRT). The hot streak temperature ratios from 1.0 (without hot streak) to 2.4 were used in these numerical simulations, including 1.0, 1.2, 1.6, 2.0, and 2.4 temperature ratios. The hot streak is circular in shape with a diameter equal to 25%of the span. The center of the hot streak is located at 50%of span and 0%of pitch (the leading edge of the HPT stator vane). The predicted results show that the hot streak is relatively unaffected as it migrates through the HPT stator. The hot streak mixes with the vane wake and convects towards the pressure surface (PS) of the HPT rotor when it moves over the vane surface of the HPT stator. The heat load of the HPT rotor increases with the increase of the hot streak temperature ratio. The existence of the inlet temperature distortion induces a thin layer of cooler air in the HPT rotor, which separates the PS of the HPT rotor from the hotter fluid. The numerical results also indicating the migration characteristics of the hot streak in the HPT rotor are predominated by the combined effects of secondary flow and buoyancy. The combined effects that induce the high-temperature fluid migrate towards the hub on the HPT rotor. The effect of the secondary flow on the hotter fluid increases as the hot streak temperature ratio is increased. The influence of buoyancy is directly proportional to the hot streak temperature ratio. The predicted results show that the increase of the hot streak temperature ratio trends to increase the relative Mach number at the HPT rotor outlet, and decrease the relative flow angle from 25%to 75%span at the HPT rotor outlet. In the other region of the HPT outlet, the relative flow angle increases when the hot streak temperature ratio is increased. The predicted results also indicate that the isentropic efficiency of the VCRT decreases with the increase of the hot streak temperature ratio.

Effect of Tip Gap and Squealer Geometry on Detailed Heat Transfer Measurements Over a High Pressure Turbine Rotor Blade Tip

The present study explores the effects of gap height and tip geometry on heat transfer distribution over the tip surface of a HPT first-stage rotor blade. The pressure ratio (inlet total pressure to exit static pressure for the cascade) used was 1.2, and the experiments were run in a blow-down test rig with a four-blade linear cascade. A transient liquid crystal technique was used to obtain the tip heat transfer distributions. Pressure measurements were made on the blade surface and on the shroud for different tip geometries and tip gaps to characterize the leakage flow and understand the heat transfer distributions. Two different tip gap-to-blade span ratios of 1% and 2.6% are investigated for a plane tip, and a deep squealer with depth-to-blade span ratio of 0.0416. For a shallow squealer with depth-to-blade span ratio of 0.0104, only 1% gap-to-span ratio is considered. The presence of the squealer alters the tip gap flow field significantly and produces lower overall heat transfer coefficients. The effects of different partial squealer arrangements are also investigated for the shallow squealer depth. These simulate partial burning off of the squealer in real turbine blades. Results show that some partial burning of squealers may be beneficial in terms of overall reduction in heat transfer coefficients over the tip surface.

Tinted soft lenses - will they replace the hard scleral lens

This paper brings out the metallurgical investigation that was carried out on discoloured low pressure turbine rotor blade (LPTR) for its platinum modified aluminium coating integrity and overheated high pressure turbine rotor blade (HPTR) for its blade material and coating integrity. LPTR blades were investigated as they were for the requirement of reuse. Heat tinting, EDS and XRD were carried out to present the harmonized inference that the coating was in good condition. HPTR blades were sectioned to study under SEM. Regions of platinum modified aluminide coating degradation and DS CM 247 LC base material cracking were observed which make the particular lot unusable.