Turbine rotor blade for gas turbine engine

Field of the Invention

0001This invention relates to a turbine rotor blade for a gas turbine engine, to a turbine rotor incorporating such blades, and to a gas turbine engine comprising such a rotor.

Background to the Invention

0002The turbine of a gas turbine engine depends for its operation on the transfer of energy between the combustion gases and turbine. The losses which prevent the turbine being totally efficient are due in part to leakage flow of working fluid over the turbine blade rotor tips.

0003In turbines with unshrouded rotor blades, a portion of the working fluid flowing through the turbine tends to migrate from the pressure surface to the suction surface of the blade aerofoil through the gap between its tip and the stationary shroud or casing. This leakage occurs because of a pressure difference that exists between the pressure and suction sides of the aerofoil. The leakage flow also causes flow disturbances to be set up over a large portion of the span of the aerofoil, which also leads to losses in the efficiency of the turbine.

0004By controlling the leakage flow of air or gas across the tips of the blades it is possible to increase the efficiency of each rotor stage. One solution is to apply a shroud to the rotor tip. When the rotor blades are assembled together in the disc that carries them, these shrouds form a continuous ring that prevents the leakage flow from the aerofoil pressure to suction side at the tip. There is still an axial leakage through the gap between the casing and the rotating shroud, but the penalties in terms of aerodynamic losses are much reduced -- often helped by the inclusion of a form of labyrinth seal on the shroud top.

0005However, the rotating shroud has a large weight penalty. As a result, the aerofoil blade speed may be constrained, to achieve acceptable blade stresses. This, though, will have the effect of increasing the aerodynamic loading that also results in reduced efficiency, negating some of the benefit of the shroud.

0006The use of a shroud ring is made more difficult if the turbine blades also operate at very high temperatures, desirable in helping to achieve high thermal efficiencies. These temperatures are limited by the turbine vane and blade materials. Cooling of these components is necessary to achieve acceptable component life, which is a function of the material temperature, stresses and material properties.

0007A large number of cooling systems are now applied to modern gas turbine blades. Such cooling systems are described for instance in <nplcit id="ncit0001" npl-type="b"><text>Cohen H, Rogers G F C, Saravanamuttoo H I H, 1981, "Gas Turbine Theory", p.232-235, Longman</text></nplcit>, and <nplcit id="ncit0002" npl-type="b"><text>Rolls-Royce pic, 1986, "The Jet Engine", p.86-88, Renault Printing Co Ltd</text></nplcit>. The more cooling that is provided, the lower the resulting material temperatures, and thus the higher the blade stresses allowable for a given component life. Cooling is achieved using relatively cool fluid bled from the upstream compressor system, bypassing the combustion chamber between the last compressor and first turbine. This air is introduced into the turbine blades where cooling is effected by a combination of internal convective cooling and external cooling. However, this cooling comes at a penalty. Its use penalises the overall efficiency of the machine, and as a result the turbine designer tries to minimise the quantity of cooling air used. All of these design constraints often leads to rotor blades of first staged turbines being shroudless -- the extra weight and thus higher stresses caused by a shroud ring simply cannot be accommodated. However, ways of reducing the high aerodynamic penalty of the resulting tip leakage flow continue to be sought.

0008<patcit id="pcit0001" dnum="US5525038A"><text>US 5,525,038</text></patcit> discloses a blade aerofoil design intended to reduce tip leakage losses. In that document, the tip region of the suction side of the aerofoil has a bowed surface with an arcuate shape. The arcuate shape of the bowed surface has progressively increasing curvature toward the tip of the rotor blade, so that a radial component of a normal to the suction side bowed surface becomes progressively larger toward the tip. It is to be noted that the aerofoil has a bowed surface in the tip region extending chordally all the way from the leading edge to the trailing edge. In addition, all the leans of the tip described, whether tangential and/or axial, are applied to the whole of the tip region.

0009A particular area of the blade that requires attention in its design is the trailing edge. Preferably, this is kept thin to minimise aerodynamic losses, but as a result it is difficult to cool, and tensile stresses have to be minimised. Cooling is achieved by films of air ejected upstream of the trailing edge on to the aerofoil surfaces, and by drilling cooling holes into the trailing edge fed from larger radial passages within the main body of the aerofoil. The aerofoil disclosed in <patcit id="pcit0002" dnum="US5525038A"><text>US 5,525,038</text></patcit> thus has some disadvantages: firstly, the curved trailing edge cannot easily have cooling holes machined into it. Ideally, this is done in one operation to minimise cost, but this requires all holes to lie in the same plane, i.e. the trailing edges of the aerofoil sections making up the blade have all to lie in one plane. A curved trailing edge (with a progressively increasing curvature) will require holes to be machined in multiple operations, incurring significant extra cost. Secondly, the leant tip will give rise to additional bending stresses in the blade. In the main body of the blades these can usually be accommodated by changes in the detailed design, such as increasing wall thickness locally. However, this cannot be done in the trailing edge region, there will simply be higher stresses in it. This will result either in a reduced life of the component, or require additional cooling, which will impair the performance of the engine.

0010The <patcit id="pcit0003" dnum="US5203676A"><text>US5203676</text></patcit> describes a tapered twisted rotating blade. The blade has a pitch to chord ratio that increases linearly with length of the blade.

0011The <patcit id="pcit0004" dnum="DE3148995"><text>DE 31 48 995</text></patcit> discloses an axial turbine with a plurality of adjustable blades

0012The <patcit id="pcit0005" dnum="JP58047102A"><text>JP 58 047102</text></patcit> shows a blade which is provided with a core to perform cooling of a rotor blade in the hollow part of the rotor blade, wherein the core is formed into shapes having twists in response to the twist in the radial direction of the rotor blade.

0013The <patcit id="pcit0006" dnum="US5352092A"><text>US 5 352 092</text></patcit> describes a turbine blade with a pressure surface, which over at least a first portion of the airfoil height has a curvature that undergoes an inflection at a location thereon, so that a first portion of the pressure surface is disposed upstream of the inflection location and a second portion of the pressure surface is disposed downstream of the inflection location. The curvature in the pressure surface first portion is opposite to the curvature in the pressure surface second portion.

Summary of the Invention

0014Accordingly, the technical problem to be solved by the present invention is to provide a turbine rotor blade for a gas turbine engine which avoids the drawbacks of the cited prior art but still provides reduced tip leakage losses. Further, the manufacturing costs should be kept low and the reliability high.

0015This problem is solved by a turbine rotor blade for a gas turbine engine having the features of claim 1. Preferred embodiments of the present invention are described in the dependent claims.

0016The invention provides a rotor blade for a gas turbine engine, the blade comprising a blade root and an aerofoil projecting therefrom, the aerofoil having a leading edge and a trailing edge, a generally concave pressure surface and a generally convex suction surface, and is characterised in that the aerofoil shape varies in section along the length thereof such that: <ul id="ul0001" list-style="dash" compact="compact"><li>the chord-wise convex curvature of the rear suction surface decreases towards the tip of the blade;</li><li>the chord-wise convex curvature of the early to mid suction surface increases towards the tip;</li><li>the stagger of the aerofoil section increases towards the tip; and</li><li>the trailing edge is a straight line.</li></ul>

0017The terms "concave" and "convex" refer in this context to the chordwise direction. Preferably, the increase in stagger is such as to give rise to chord-wise convex curvature on the pressure surface. This has the aerodynamic result that velocities on the pressure surface are increased, reducing the local static pressure which drives the over-tip leakage, thus reducing the leakage further.

0018The line of the trailing edge is advantageously radial, but it may be leant at some radial and/or tangential angle, depending on the aerodynamic and mechanical details of the design.

0019The chord-wise convex curvature of the rear suction surface may be eliminated in the tip section to give a flat back surface, or it may even be reversed so as to be concave.

0020The lower part of the aerofoil will preferably have a minimum radial extent of 40% and a maximum radial extent of 80% of the span. The tip section is preferably geometrically and aerodynamically blended smoothly into the lower portion of the aerofoil. The blending of the aerofoil between the tip section and the lower portion will preferably be such as to evenly redistribute the aerodynamic loading radially away from the tip.

0021The trailing edge regions of the aerofoil sections are preferably shaped so as to be similar in 2-D plan view, such that a row of chord-wise running cooling holes provided in the trailing edge will lie in a radial geometric plane, enabling them to be machined in a single operation. The increased stagger of the aerofoil tip section, relative to the lower, datum portion, will cause the front/mid portion of the aerofoil tip region to be tangentially leant, in a direction towards the suction side. Unlike the arrangements disclosed in <patcit id="pcit0007" dnum="US5525038A"><text>US 5,525,038</text></patcit>, the leaning is confined to a limited part of the tip, while the trailing edge is kept straight. Thus, in the tip region only the front/mid portion of the suction surface may be concave in the radial direction, and only the front portion of the pressure surface may be convex in the radial direction.

0022The invention also provides a turbine rotor for a gas turbine engine, comprising a plurality of turbine blades in accordance with the invention. The invention further provides a gas turbine engine including the turbine rotor of the invention.

Brief Description of the Drawings

0023A more complete appreciation of the invention and of the attendant advantages thereof will be readily obtained as the same becomes better understood by reference to the following detailed description when considered in connection with the accompanying drawings. In the drawings: <ul id="ul0002" list-style="none"><li><figref idref="f0001">Figure 1</figref> is a plot of Mach number distributions (normalised to exit) for the mid-span aerofoil sections of two rotor blades according to the prior art;</li><li><figref idref="f0001">Figure 2</figref> is an illustration of the mid-span section profile shapes for the blades plotted in <figref idref="f0001">Figure 1</figref>;</li><li><figref idref="f0002">Figures 3 and 4</figref> are isometric views of the full aerofoil shapes of the blades of <figref idref="f0001">Figures 1 and 2</figref>;</li><li><figref idref="f0003">Figure 5</figref> is a plot of percentage loss of turbine stage efficiency against tip clearance or gap as a percentage of blade span, produced by Computational Fluid Dynamics (CFD);</li><li><figref idref="f0003">Figure 6</figref> is a plot of tip Mach number distribution, as calculated by CFD, for the aerofoils shown in <figref idref="f0002">Figures 3 and 4</figref>;</li><li><figref idref="f0004">Figure 7</figref> is an isometric view of two high lift aerofoils which have been restacked to give a lean;</li><li><figref idref="f0004">Figure 8</figref> repeats the plot of <figref idref="f0003">Figure 5</figref>, but with comparative results for the aerofoil of <figref idref="f0004">Figure 7</figref> included;</li><li><figref idref="f0005">Figure 9</figref> is a plot of the calculated Mach number distribution for the aerofoils of <figref idref="f0002">Figures 3</figref> and <figref idref="f0004">7</figref> at 90% of the span;</li><li><figref idref="f0005">Figure 10</figref> is a corresponding plot at mid-span;</li><li><figref idref="f0006">Figure 11</figref> is a plot of calculated 2-D Mach number distributions for the conventional and redesigned blade illustrated in <figref idref="f0006">Figure 12</figref>;</li><li><figref idref="f0006">Figure 12</figref> is a sectional view comparing a conventional aft-loaded aerofoil tip section with a corresponding section of a blade in accordance with the invention;</li><li><figref idref="f0007">Figure 13</figref> is an isometric view from the front of rotor blades in accordance with a preferred embodiment of the invention;</li><li><figref idref="f0007">Figure 14</figref> is a plot of tip Mach number distribution, as calculated by CFD, for the aerofoil in accordance with the invention;</li><li><figref idref="f0008">Figure 15</figref> is a comparison of CFD calculated rotor loss against tip gap; and</li><li><figref idref="f0008">Figure 16</figref> is a plan view of the rotor blade of <figref idref="f0007">Figure 13</figref> showing successive aerofoil sections equally-spaced radially at 5% of span.</li></ul>

Detailed Description of the Illustrated Embodiment

0024Before describing in detail the geometric configuration of a turbine rotor blade in accordance with the invention, it is useful to describe how the configuration operates, and in particular how it affects the turbine rotor aerodynamics. First the basic turbine aerodynamics will be considered, then the benefits of aerodynamically off-loading the tip, and finally the detail of the invention will be described.

0025The turbine blade of the invention is of a "high lift" aerodynamic design, and <figref idref="f0001">Figure 1</figref> is a graph comparing the Mach number distributions (normalised to exit) for the mid-span aerofoil sections of two rotor blades that have the same axial chord and the same inlet and exit flow conditions as each other. The two aerofoils are differentiated as follows: <ol id="ol0001" compact="compact"><li>a) A conventional (low) lift aerofoil characterised by only a small diffusion of the flow from the Mach number point on the late suction surface to the trailing edge (known as " back surface diffusion").</li><li>b) A high lift aerofoil, carrying approximately 36% more lift than a conventional profile, achieved by increasing the pitch of the profiles in the same proportion to give a reduction in aerofoil numbers of about 30%. The back surface diffusion is now much larger, and the peak Mach number has significantly increased.</li></ol>

0026<figref idref="f0001">Figure 2</figref> compares the mid-span section profile shapes. The full aerofoil shapes of the two blades are shown in isometric view in <figref idref="f0002">Figure 3</figref> (conventional) and <figref idref="f0002">Figure 4</figref> (high lift). The increased pitch of the high lift blading is clear. Both aerofoils clearly have convex curved suction surfaces 1 (in the chord-wise direction), with the locations of the peak Mach number point coinciding with a local maximum in the surface curvature. Each blade has concave pressure surfaces 2, a root end 3 and a tip 4. Both aerofoils are radially stacked in the same way on a straight line 5 through the centres of the trailing edge circles.

0027Calculations of the flow around these two aerofoils have been performed using Computational Fluid Dynamics (CFD) over a range of tip clearances. The results, in the form of predicted rotor loss (expressed as a percentage of turbine stage efficiency) against tip gap (expressed as a percentage of rotor span) are shown in <figref idref="f0003">Figure 5</figref>. It should be noted that these predictions would not normally be used to give absolute values of efficiency, but may still demonstrate qualitative differences between different geometries.

0028<nplcit id="ncit0003" npl-type="s"><text>ASME 96-TA-13, November 1998, "Reduction of Tip Clearance Losses Through 3-D Airfoil Designs", Staubach J B, Sharma O P, Stetson G M</text></nplcit> (reference 3) gives a typical value for the over-tip leakage loss exchange rate of a turbine rotor blade of about numeral 2% of stage efficiency for 1% gap/span. The values shown in <figref idref="f0003">Figure 5</figref> for the conventional blade are slightly less than this, confirming that CFD is best used for qualitative predictions, rather than absolute accuracy.

0029<figref idref="f0003">Figure 5</figref> also shows that at zero tip gap, the two aerofoils have almost the same loss (given the accuracy of the CFD code used). However, it is clear that the high lift aerofoil has significantly higher tip leakage losses than the conventional one, at any given tip gap. The reason for this can be seen in <figref idref="f0003">Figure 6</figref>. This compares the calculated Mach number distributions at 90% span, near the tip, for the conventional and high lift aerofoils respectively. The much higher aerodynamic loading of the high lift aerofoil gives rise to higher tip leakage and higher mixing losses, due to the higher gas stream velocities on the suction side.

0030Reference 3 shows how leaning the blade tip, as illustrated in <patcit id="pcit0008" dnum="US5525038A"><text>US 5,525,038</text></patcit>, reduces the tip leakage loss. To demonstrate this, the high lift aerofoil has been stacked as described in Reference 3, and the resulting flow field calculated again by CFD. <figref idref="f0004">Figure 7</figref> shows an isometric view of the restacked, high lift aerofoil. The lean is purely tangential, beginning at about 60% span, the curve of the stacking axis being parabolic in shape with the highest angle to the vertical (40°) being at the tip.

0031<figref idref="f0004">Figure 8</figref> repeats the plot of <figref idref="f0003">Figure 5</figref> (rotor loss as a percentage of the stage efficiency versus percentage gap/span) but now with the leant rotor tip results added. The tip leakage loss is reduced, but the zero gap loss is higher, which means that it is only really beneficial at larger tip gaps.

0032The reasons for this can be understood by comparing the calculated aerofoil Mach number distributions for the straight and leant tip aerofoils at 90% span (<figref idref="f0005">Figure 9</figref>), and at mid-span (<figref idref="f0005">Figure 10). Figure 9</figref> shows how the tip lean has off-loaded the need suction surface of the tip section; this results in reduced tip leakage loss. However, the loading has been redistributed down the span, as can be seen at mid-span in <figref idref="f0005">Figure 10</figref>. The higher surface velocities here result in increased profile (wetted area) losses, and although these are somewhat reduced near the tip, the overall effect is an increase in loss at zero gap.

0033<figref idref="f0005">Figure 9</figref> is also noteworthy in that the tip lean is shown to only off-load the early/mid part of the aerofoil suction surface. This is one of the drawbacks of simply leaning the aerofoil tip whilst leaving the aerofoil shapes unchanged. The velocity distribution on the late suction surface is largely unaffected, which limits the reduction in the aerodynamic loading at the tip that can be achieved.

0034Referring now to <figref idref="f0006">Figure 12</figref>, in accordance with the invention, the 2-D aerofoil section at the tip is redesigned to significantly change the velocity distribution. Instead of being aft-loaded, the loading is moved forward, when analysed as simply a 2-D aerodynamic section. This can be seen in <figref idref="f0006">Figure 11</figref>, which compares the calculated 2-D Mach number distributions for the original and redesigned tip sections at 90 percent span. The redesigned profile shown is an extreme example, where the loading has been moved to the front of the aerofoil. More usually, the loading will be moved to the mid region. The design of the aerofoils in the lower portion of the blade, radially from the hub to, typically, around mid-span remains strongly aft-loaded. The change in the 2-D lift distribution of the tip section is effected by reducing the chord-wise convex curvature on the late suction surface and increasing this curvature in the front/mid region. It is possible that the late suction surface can become flat or even concave, to minimise the lift locally. As a result of the changes in the surface curvature of the tip section, the stagger of the section increases. Stagger is here defined as the angle between the turbine centre line and a line drawn through the centres of the leading and trailing edge circles of the aerofoil. Thus, the front part of the redesigned aerofoil is moved in the direction of the suction surface, relative to the original design, when viewed with the trailing edges coincident, as shown in <figref idref="f0006">Figure 12</figref>.

0035The redesigned tip section, with lift moved forward, is now stacked with the aft-loaded aerofoils in the lower portion to produce the aerofoil shown in isometric view in <figref idref="f0007">Figure 13</figref>. The stack is on a radial line through the trailing edge, as with the original high lift aerofoil. It can be seen that with the radially straight trailing edge 5, the progressive increase in aerofoil stagger up the span results in local tip lean in the forward part of the aerofoil. This local tip lean has the effect of redistributing the aerodynamic loading in the front part of the aerofoil radially, from the tip downwards to the lower aerofoil sections. It should be remembered that the 2-D design of the tip section deliberately moved the lift forward to the mid/front regions of the aerofoil, where the local lean has the most effect. The redistribution of lift radially inwards from the tip region is similar to that caused by leaning the whole tip, and results in increased lift on the lower aerofoil sections. Since this extra lift is in the front/mid part of the suction surface, it can be seen that designing these lower sections to be strongly aft-loaded will in part compensate for this.

0036The resulting tip Mach number distribution, as calculated by CFD, is shown in <figref idref="f0007">Figure 14</figref> compared with that of the original high lift design. It will be seen that the Mach numbers along most of the suction surface have been reduced, compared with the unmodified aerofoil. This will result in lower tip leakage flow and reduced mixing. The Mach numbers on the late pressure surface have been raised, that is the local static pressures have been reduced. This will also have the effect of reducing tip leakage. There is a small increase in lift on the suction surface, just after the leading edge, for the redesigned profile. This is because an extreme forward loaded design has been shown as an example here. However, this is not an intrinsic feature of the invention. Comparison of <figref idref="f0006">Figure 11</figref> with <figref idref="f0007">Figure 14</figref> shows that the high lift on the early suction surface in the 2-D design has been removed (radially redistributed) by 3-D effects, as intended.

0037<figref idref="f0008">Figure 15</figref> plots the variation of CFD predicted rotor loss with tip gap, repeating the results of <figref idref="f0004">Figure 8</figref>, but now with the results from the redesigned tip added. It can be seen from the Figure that at zero tip gap, the loss is higher than that of the unmodified high lift aerofoil, and very similar to that of the fully leant tip. The reasons for the loss being higher than the original profile at zero tip gap are much the same as for the fully leant tip. The loss due to over tip leakage is much reduced compare to the unmodified high lift aerofoil. The improvement is better than the fully leant aerofoil analysed here. Further, for large gaps (two percent of span), the loss is close to that of the low lift aerofoil. The increased loss due to high lift has almost been eliminated.

0038<figref idref="f0008">Figure 16</figref> shows an overlay of some of the 2-D aerofoil sections, viewed radially from above. It can be seen that, in the trailing edge region, the aerofoil shapes are almost coincident. This means that machining of trailing edge cooling holes can easily be done in a single operation, minimising cost of manufacture, and there should be no additional bending stresses in this critical trailing edge region.

Reference Signs

0039<dl id="dl0001" compact="compact"><dt>1</dt><dd>suction surface</dd><dt>2</dt><dd>pressure surface</dd><dt>3</dt><dd>blade root</dd><dt>4</dt><dd>tip</dd><dt>5</dt><dd>trailing edge</dd></dl>