Showing papers by "Rolls-Royce Holdings published in 2001"

Limitations on Gas Turbine Performance Imposed by Large Turbine Cooling Flows

Abstract: Calculations of the performance of modern gas turbines usually include allowance for cooling air flow rate; assumptions are made for the amount of the cooling air bled from the compressor, as a fraction of the mainstream flow, but this fractional figure is often set in relatively arbitrary fashion. There are two essential effects of turbine blade cooling: (i) the reduction of the gas stagnation temperature at exit from the combustion chamber (entry to the first nozzle row) to a lower stagnation temperature at entry to the first rotor and (ii) a pressure loss resulting from mixing the cooling air with the mainstream, Similar effects occur in the following cooled blade rows. The paper reviews established methods for determining the amount of cooling air required and semi-empirical relations, for film cooled blading with thermal barrier coatings, are derived. Similarly, the pressure losses related to elements of cooling air leaving at various points round the blade surface are integrated over the whole blade. This gives another semi-empirical expression, this time for the complete mixing pressure loss in the blade row, as a function of the total cooling air used. These two relationships are then used in comprehensive calculations of the performance of a simple open-cycle gas turbine, for varying combustion temperature and pressure ratio. These calculations suggest that for maximum plant efficiency there may be a limiting combustion temperature (below that which would be set by stoichiometric combustion). For a given combustion temperature, the optimum pressure ratio is reduced by the effect of cooling air.

A converging slot-hole film-cooling geometry - Part 1: Low-speed flat-plate heat transfer and loss

Abstract: This paper presents experimental measurements of the performance of a new film cooling hole geometry - the Converging Slot-Hole or Console. This novel, patented geometry has been designed to improve the heat transfer and aerodynamic loss performance of turbine vane and rotor blade cooling systems. The physical principles embodied in the new hole design are described, and a typical example of the console geometry is presented. The cooling performance of a single row of consoles was compared experimentally with that of typical 35° cylindrical and fan-shaped holes and a slot, on a large-scale, flat-plate model at engine representative Reynolds numbers in a low speed tunnel with ambient temperature main flow. The hole throat area per unit width is matched for all four hole geometries. By independently varying the temperature of the heated coolant and the heat flux from an electrically heated, thermally insulated, constant heat flux surface, both the heat transfer coefficient and the adiabatic cooling effectiveness were deduced from digital photographs of the colour play of narrowband thermochromic liquid crystals on the model surface. A comparative measurement of the aerodynamic losses associated with each of the four film-cooling geometries was made by traversing the boundary layer at the downstream end of the flat plate. The promising heat transfer and aerodynamic performance of the console geometry have justified further experiments on an engine representative nozzle guide vane in a transonic annular cascade presented in Part 2 of this paper [1].

A Converging Slot-Hole Film-Cooling Geometry—Part 2: Transonic Nozzle Guide Vane Heat Transfer and Loss

Abstract: This paper presents the first experimental measurements on an engine representative nozzle guide vane, of a new film-cooling hole geometry, a convergingslot-hole or console. The patented console geometry is designed to improve the heat transfer and aerodynamic performance of turbine vane and rotor blade cooling systems. These experiments follow the successful validation of the console design in low-speed flat-plate tests described in Part 1 of this paper. Stereolithography was used to manufacture a resin model of a transonic, engine representative nozzle guide vane in which seven rows of previously tested fan-shaped film-cooling holes were replaced by four rows of consoles. This vane was mounted in the annular vane ring of the Oxford cold heat transfer tunnel for testing at engine Reynolds numbers, Mach numbers and coolant to mainstream momentum flux ratios using a heavy gas to simulate the correct coolant to mainstream density ratio. Heat transfer data were measured using wide-band thermochromic liquid crystals and a modified analysis technique. Both surface heat transfer coefficient and the adiabatic cooling effectiveness were derived from computer-video records of hue changes during the transient tunnel run. The cooling performance, quantified by the heat flux at engine temperature levels, of the console vane compares favourably with that of the previously tested vane with fan-shaped holes. The new console film-cooling hole geometry offers advantages to the engine designer due to a superior aerodynamic efficiency over the fan-shaped hole geometry. These efficiency measurements are demonstrated by results from midspan traverses of a four-hole pyramid probe downstream of the nozzle guide vane.

Improving the Efficiency of the Trent 500 HP Turbine Using Non-Axisymmetric End Walls: Part 1 — Turbine Design

Abstract: This paper describes the redesign of the HP turbine of the Rolls-Royce Trent 500 engine, making use of non-axisymmetric end walls. The original, datum turbine used conventional axisymmetric end walls, while the vane and (shrouded) rotor aerofoil profiles were nominally the same for the two designs. Previous research on the large scale, low speed linear cascade at Durham University, see Hartland et al [1], had already demonstrated significant potential for reducing turbine secondary losses using non-axisymmetric end walls - by about one third. This paper shows how a methodology was derived from the results of this research and applied to the design of the single stage Trent 500 HP turbine (model rig). In particular the application of a new linear design system for the parametric definition of these end wall shapes, described in Harvey et al [2], is discussed in detail.Copyright © 2001 by ASME

The Application of Ultra High Lift Blading in the BR715 LP Turbine

Abstract: The original LP turbine of the BR715 engine featured “High Lift” blading, which achieved a 20% reduction in aerofoil numbers compared to blading with conventional levels of lift - reported in Cobley et al. (1997). This paper describes the design and test of a re-bladed LP turbine with new “Ultra High Lift” aerofoils, achieving a further reduction of approximately 11% in aerofoil count and significant reductions in turbine weight. The design is based on the successful cascade experiments of Howell et al. (2000) and Brunner et al. (2000). Unsteady wake - boundary layer interaction on these low Reynolds number aerofoils is of particular importance in their successful application. Test results show the LP turbine performance to be in line with expectation. Measured aerofoil pressure distributions are presented and compared with the design intent. Changes in the turbine characteristics relative to the original design are interpreted by making reference to the detailed differences in the two aerofoil design styles.Copyright © 2001 by ASME

Mode detection with an optimised array in a model turbofan engine intake at varying shaft speeds

Abstract: Modal measurement techniques in engine intakes have been used previously to analyse the generated fan noise. A proven method is to use a wall-mounted array of Kulite transducers and operate the (model) turbofan under constant shaft speeds. A drawback of this method is the large number of (expensive) microphones and acquisition channels needed to obtain complete m-mode spectra at high engine orders. Furthermore, to get a full scan of the m-mode spectra as a function of shaft speed, many measurements are required. The issue of the large number of microphones was addressed by using a sparse array instead of an equidistant array. An array optimisation technique, similar to a technique used for the design of phased microphone arrays for sound source localisation, was used to define such a sparse intake array. This array consists of 100 Kulites and is able to determine without aliasing the modal spectrum from m = -79 to m = -1-79, which is appropriate to determine the modal content up to 3 BPF of a modem turbofan. This array was tested in a Rolls-Royce model fan rig at Ansty as a part of the RESOUND project. A new digital data-acquisition system made it possible to simultaneously and continuously record the Kulite pressure data as the engine speed was varied continuously from idle to maximum speed or vice versa, with each acceleration/deceleration lasting for a period of 9 minutes. Time histories of the Kulites were processed giving power spectra of the engine orders, which revealed the rotor locked tonal components. For each rotor revolution, a Discrete Fourier Transform was applied and, after averaging over a number of revolutions, the m-mode spectra were determined. In this way, a full modal scan with respect to shaft speed in a very limited testing time was obtained. Paper presented at the 7th AIAA/CEAS Aeroacoustics Conference, 28-30 May 2001, Maastricht, The Netherlands.

Improving the Efficiency of the Trent 500 HP Turbine Using Non-Axisymmetric End Walls: Part II — Experimental Validation

Abstract: Part I of this paper described how the HP turbine model rig of the Rolls-Royce Trent 500 was redesigned by applying non-axisymmetric end walls to both the vane and blade passages, whilst leaving the turbine operating point and overall flow conditions unaltered. This paper describes the results obtained from testing of the model rig and compares them with those obtained for the datum design (with conventional axisymmetric end walls). Measured improvements in the turbine efficiency are shown to be in line with those expected from the previous linear cascade research at Durham University, see Harvey et al. [1] and Hartland et al. [2]. These improvements are observed at both design and off-design conditions. Hot wire traverses taken at the exit of the rotor show, unexpectedly, that the end wall profiling has caused changes across the whole of the turbine flow field. This result is discussed making reference to a preliminary 3-D CFD analysis. It is concluded that the design methodology described in part I of this paper has been validated, and that non-axisymmetric end wall profiling is now a major new tool for the reduction of secondary loss in turbines (and potentially all axial flow turbomachinery). Further work, though, is needed to fully understand the stage (and multistage) effects of end wall profiling.Copyright © 2001 by ASME

Spatter-free laser percussion drilling of closely spaced array holes

Abstract: Spatter is one of the inherent defects commonly associated with holes produced with laser drilling. This work reports on a method of spatter prevention, workable for a wide range of process parameters. The method is based on the application of a specially developed anti-spatter composite coating (ASCC), containing a mixture of ceramic filler particles embedded in a silicone elastomer matrix, on the workpiece surface prior to laser percussion drilling. Experiments were conducted using a fibre-optic delivered 400 W Nd:YAG laser for the drilling of closely spaced through holes (2 mm hole pitch) in Nimonic 263 alloy sheets. The work revealed that the ASCC effectively prevented the deposition of spatter such that laser drilled holes were produced whilst maintaining the as-received surface characteristics of the Nimonic 263 alloy for all the assist gases tested (O 2 , air, N 2 and Ar). The process characteristics and spatter prevention mechanism associated with the use of the ASCC have been investigated using optical microscopy, scanning electron microscopy (SEM), energy dispersive X-ray (EDX) analysis, contact angle analysis and high-speed photographic imaging. Comparative studies were also made with the case of uncoated laser drilling.

Boundary Layer Development in the BR710 and BR715 LP Turbines: The Implementation of High Lift and Ultra High Lift Concepts

Abstract: This paper describes a detailed study into the unsteady boundary layer behaviour in two high lift and one ultra high lift Rolls-Royce Deutschland LP turbines. The objectives of the paper are to show that high lift and ultra high-lift concepts have been successfully incorporated into the design of these new LP turbine profiles. Measurements from surface mounted hot film sensors were made in full size, cold flow test rigs at the altitude test facility at Stuttgart University. The LP turbine blade profiles are thought to be state of the art in terms of their lift and design philosophy. The two high lift profiles represent slightly different styles of velocity distribution. The first high-lift profile comes from a two stage LP turbine (the BR710 cold-flow, high-lift demonstrator rig). The second high-lift profile tested is from a three-stage machine (the BR715 LPT rig). The ultra-high lift profile measurements come from a redesign of the BR715 LP turbine: this is designated the BR715UHL LP turbine. This ultra high-lift profile represents a 12% reduction in blade numbers compared to the original BR715 turbine. The results from NGV2 on all of the turbines show “classical” unsteady boundary layer behaviour. The measurements from NGV3 (of both the BR715 and BR715UHL turbines) are more complicated, but can still be broken down into classical regions of wake-induced transition, natural transition and calming. The wakes from both upstream rotors and NGVs interact in a complicated manner, affecting the suction surface boundary layer of NGV3. This has important implications for the prediction of the flows on blade rows in multistage environments. NOMENCLATURE w τ Quasi wall shear stress w τ ′ RMS of signal w τ ~ Non-dimensional ensemble mean quasi wall shear stress

Device and method for fatigue testing of materials

Abstract: A device (10) for fatigue testing of materials comprises a frame (14), first and second clamping means (16,18) for holding a specimen (12) to be tested. First and second mounting means (20,22) mount the clamping means (16,18) on the frame (14). The mounting means (20,22) vibrationally isolate the clamping means (16,18) from the frame (14). Actuator means (24) moves the first clamping means (16) relative to the second clamping means (18) to apply a low cycle load on the specimen (12). Electrical insulating means (30) electrically insulate the frame (14) from the specimen (12). A shaker (26) is coupled to the second clamping means (18) to apply a high cycle load on the specimen (12). A detector (32) detects the vibration of the specimen (12) and sends an electrical signal to a control unit (42) which determines the resonant frequency of the specimen (12). The control unit (42) sends a signal to the shaker (26) to maintain the high cycle load at the resonant frequency of the specimen (12). Electrical potential drop probes (38) are provided on the specimen (12) to send a second electrical signal to the control unit (42) which is arranged to determine the rate of crack growth, the fatigue life to crack initiation and fatigue life to failure of the specimen (12).

The role of customer relationships in the growth of small‐ to medium‐sized manufacturers

Abstract: This work is focused on small‐ to medium‐sized manufacturers (SMMs) within supply chains and, in particular, on how such SMMs develop effective working relationships with customers. Development of a model based on the literature was followed by a mail survey, augmented by semi‐structured interviews with SMMs. Factor analysis and multiple regression analysis were used to provide an understanding of the underlying processes.

Mechanisms and Prediction Methods for Fan Blade Stall Flutter

Abstract: A e rst aim is to distinguish between two e utter regimes that are commonly referred to as stall e utter because of their occurrence between a raised working line and the stall boundary. The e rst one is of aeroacoustic origin, and it arises from a match between the acoustic impedance of the intake duct and the upstream pressure perturbation due to fan vibration. The second type is directly related to e ow separation effects and shock properties. Further objectives are to describe several numerical prediction methods, to compare their relative computational requirements, and to establish their bounds of applicability to various e utter types. The unsteady e ow is either a linearization about a viscous nonlinear steady-state e ow for a given mode of vibration, or it is fully nonlinear. The prediction methods are classie ed according to the way they treat the unsteady e ow. In all cases, the e uid mesh was moved during the unsteady e ow computations to follow the structural motion. The performance of the methods was ranked for a rig fan blade for which measurements were available. It was found that, near the stall boundary,e utterboundary couldonlybecaptured withan adequaterepresentationoftheunsteady viscouseffects. It was concluded that the shock had a stabilizing effect whereas the separation area behind it had a destabilizing effect.

Method of manufacturing an article by diffusion bonding and superplastic forming

Abstract: A method of manufacturing a gas turbine engine fan blade ( 10 ) comprises forming three metal workpieces ( 30, 32, 34 ). The metal workpieces ( 30, 32, 34 ) are assembled into a stack ( 36 ) so that the flat surfaces ( 38, 42, 46, 48 ) are in mating abutment. Heat and pressure is applied across the thickness of the metal workpieces ( 30, 32, 34 ) to diffusion bond the metal workpieces ( 30, 32, 34 ) together to form an integral structure ( 80 ). The integral structure ( 80 ) is upset forged at one end ( 58 ) to produce an increase in thickness ( 82 ) for forming the blade root ( 26 ). The upset forged integral structure ( 80 ) is then hot creep formed and superplastically formed to produce the required aerofoil shape and the thickened end ( 82 ) is machined to form the blade root ( 26 ). The method enables thinner metallic workpieces with better microstructure to be used and increases the yield of metallic workpieces.

Cooling of gas turbine engine aerofoils

Abstract: An aerofoil ( 39 ) for a gas turbine engine ( 10 ) includes one or more internal cooling passage ( 74,76 ) for receiving a flow of cooling fluid, and means ( 84,86,88,90 ) for inducing at least two vortices in cooling fluid flowing through each cooling passage. Fluid within the vortices has a radially outwards, screw-type motion. The two vortices produce effective cooling of the aerofoil, which may be an aerofoil of a turbine blade.

Coupled fluid/solid heat transfer computation for turbine discs

Abstract: This paper considers the coupling of a finite element thermal conduction solver with a steady, finite volume fluid flow solver. Two methods were considered for passing boundary conditions between the two codes - transfer of metal temperatures and either convective heat fluxes or heat transfer coefficients and air temperatures. These methods have been tested on two simple rotating cavity test cases and also on a more complex real engine example. Convergence rates of the two coupling methods were compared. Passing heat transfer coefficients and air temperatures was found to give the quickest convergence. The coupled method gave agreement with the analytic solution and a conjugate solution of the simple free disc problem. The predicted heat transfer results for the real engine example showed some encouraging agreement, although some modelling issues are identified. Copyright © 2001 by ASME.

Low Thermal Conductivity EB-PVD Thermal Barrier Coatings

Abstract: The continued drive for increased performance and higher engine efficiencies has seen the adoption of thermal barrier coating technologies as an integral part of modern engine design. Future engines will require even better performance from such thermal barrier coating system, towards the ultimate goal of designed-in TBC's. This paper reviews the advantages and disadvantages of various TBC systems, with the aim of custom designing a thermal barrier coating system to have good strain tolerance and low thermal conductivity. The influence of coating microstructure on the coating thermal conductivity and strain tolerance is discussed, and it is shown that by advanced processing it is possible to deposit EB-PVD coatings with thermal conductivities of 1.0W/mK, much lower than current commercial EB-PVD coatings at 1.5-1.9 W/mK and closely matching that for plasma sprayed ceramics (0.8-1.1 W/mK).

A combined testing and modelling approach to the prediction of the fretting fatigue performance of splined shafts

Abstract: Splined shafts used in gas turbines are being subjected to an ever more arduous set of loads, but the strength of this class of joints has hitherto not been probed. This paper describes preliminary contributions to determining this by first performing a finite element analysis of the spline overall in order to determine the contact conditions, including surface slip displacement. These conditions are then reproduced in a simplified fretting fatigue apparatus in order to reduce the number of prototypical tests required.

Friction Damper Optimization: Simulation of Rainbow Tests

Abstract: Friction dampers have been used to reduce turbine blade vibration levels for a considerable period of time. However, optimal design of these dampers has been quite difficult due both to a lack of adequate theoretical predictions and to difficulties in conducting reliable experiments. One of the difficulties of damper weight optimization via the experimental route has been the inevitable effects of mistuning. Also, conducting separate experiments for different damper weights involves excessive cost. Therefore, current practice in the turbomachinery industry has been to conduct so-called rainbow tests where friction dampers with different weights are placed between blades with a predefined configuration. However, it has been observed that some rainbow test results have been difficult to interpret and have been inconclusive for determining the optimum damper weight for a given bladed-disk assembly. A new method of analysis-a combination of the harmonic balance method and structural modification approaches-is presented in this paper for the analysis of structures with friction interfaces and the method is applied to search for qualitative answers about the so-called rainbow tests in turbomachinery applications. A simple lumped-parameter model of a bladed-disk model was used and different damper weights were modeled using friction elements with different characteristics. Resonance response levels were obtained for bladed disks with various numbers of blades under various engine-order excitations. It was found that rainbow tests, where friction dampers with different weights are used on the same bladed-disk assembly, can be used to find the optimum damper weight if the mode of vibration concerned has weak blade-to-blade coupling (the case where the disk is almost rigid and blades vibrate almost independently from each other). Otherwise, it is very difficult to draw any reliable conclusion from such expensive experiments.

Non-axisymmetric turbine end wall profiling:

Abstract: A design method for profiling the end wall to reduce secondary flow has been reported previously. A profile has been tested in the Durham Linear Cascade and the results confirmed the design method. This paper describes the design and testing of a second-generation end wall, where the profiling is more suited to a real turbine. The new end wall has been tested in the linear cascade and a comprehensive set of measurements have been taken. These include traverses of the flow field upstream and downstream of the blade row, surface static pressure distributions on the end wall and flow visualization. Comparisons have been made with the results with a planar end wall and the earlier profiled end wall. Observed reductions in exit angle deviations are even greater than for the first design, although the loss reduction is not as great. The results verify the design, confirming profiled end walls as a means of reducing secondary flow, kinetic energy and loss. Overall an improved understanding of the effects of end wall profiling has been obtained although further work is required in this area.

Seal segment for a turbine

Abstract: A seal segment ( 66 ) as described for a seal segment ring ( 64 ) of a turbine ( 16 ) in a gas turbine engine ( 10 ). The seal segment ( 66 ) has an inner surface ( 70 ) adapted to face the turbine blades ( 36 ) in use. Path means ( 72 ) is defined in the seal segment ( 66 ). The path means ( 72 ) is adapted to extend, in use, generally parallel to the principal axis of the turbine and has downstream inlet means ( 74 ) through which a cooling fluid to cool the seal segment can enter the path means ( 72 ) and upstream outlet means ( 76 ) from which the cooling fluid can be exhausted from the path means ( 72 ). The cooling fluid can flow along the path means ( 72 ) in a generally upstream direction opposite to the flow of gas through the turbine.

Gas turbine engine rotor blades

Abstract: A component for a gas turbine engine having at least one surface, that has been treated by ultrasonic hammer peening so as to provide a region of deep compressive residual stress in the treated region.