Patent Publication Number: US-8523523-B2

Title: Cooling arrangements

Description:
BACKGROUND 
     The present invention relates to cooling arrangements and more particularly to cooling arrangements in blades such as high pressure turbine blades in a gas turbine engine. 
     With high pressure turbine blades within gas turbine engines it will be appreciated that the relatively high temperatures to which the blades are subjected necessitate cooling in order that the materials from which such components are made can remain within the operational capabilities of those materials. Other components within a gas turbine engine which must be able to withstand such high temperatures and other operational requirements include nozzle guide vanes. Traditionally two approaches have been taken with regard to achieving necessary cooling. Firstly, impingement cooling is achieved through providing passages which extend along the length of the blade or other component with a coolant fluid under pressure, which then is projected through impingement orifices from the passage to a chamber beneath the surface to be cooled. In such circumstances, coolant fluid is projected towards that surface at high velocity, generating high heat transfer, thereby coking that part of the component. An alternative is simply provision of radial channels which are presented below the surface of the component. Each approach has its advantages and disadvantages. Impingement cooling generally gives significantly increased heat transfer compared to radial cooling even where ribs are utilised to create turbulence, but the necessity for impingement orifices greatly increases manufacturing complexity, cost and may reduce fatigue life. 
     SUMMARY 
     It will be appreciated that the leading edge of a turbine blade has a high external heat flux and in such circumstances requires significant amounts of film cooling to protect against oxidation and fatigue damage. Furthermore in situations where a thermal barrier coating is used such locations are also vulnerable to the coating being lost through foreign object damage or over temperature of the coating and/or its bond coat which can further shorten operational life. Through use of appropriate cooling technology, improvements can be made which reduce the leading edge temperature, but a balance must be struck between reducing cooling air consumption and allowing an increase in the temperature at which the engine operates which in turn will affect overall engine performance in terms of efficiency and reduced fuel burn. 
     According to aspects of the present invention there is provided a cooling arrangement for a hollow blade, the arrangement comprising a passage for a fluid flow therealong, opposed undulations provided in the passage to engage the fluid flow in use to generate a lateral or rotating vortex flow aspect in the fluid flow and a shaped portion of the passage between the opposed undulations shaped to divide the vortex flow aspect into a number of vortices. 
     Typically, the shaped portion of the passage is angular. Generally, the undulations are ribs or turbulators. 
     Possibly, the shaped portion includes undulations to facilitate vortex development. 
     Possibly, the passage has an adjacent wall containing impingement orifices opposite the shaped portion, these impingement orifices connect to a further passage. Typically, the orifice portion is also shaped to facilitate vortex development in the passage. 
     Possibly, the orifice portion divides the passage from a leading passage in a hollow blade. 
     Generally, the orifices of the orifice portion are directed to project at least a proportion of the fluid flow towards an opposed portion of the leading passage. 
     Generally, the shaped portion is arranged in the passage whereby the vortices are substantially constrained within their respective portion of the passage. 
     Also in accordance with aspects of the present invention there is provided a blade incorporating a cooling arrangement as described above. Typically, the blade is a high pressure turbine blade for a gas turbine engine. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
       Embodiments of aspects of the present invention will now be described by way of example only with reference to the accompanying drawings in which: 
         FIG. 1  is a schematic section through a conventional gas turbine engine in which a blade in accordance with the present invention may be used; 
         FIG. 2  is a schematic cross section of a typical prior cooling arrangement; 
         FIG. 3  provides a schematic cross section of a first embodiment of aspects of the present invention; 
         FIG. 4  provides a schematic illustration of a variant of the first embodiment of aspects of the present invention as depicted in  FIG. 2  in greater detail; 
         FIG. 5  is a schematic illustration of a second embodiment of aspects of the present invention; 
         FIG. 6  is a schematic cross section of a third embodiment of aspects of the present invention; 
         FIG. 7  is a schematic cross section of a fourth embodiment of aspects of the present invention; 
         FIG. 8  is a schematic cross section of a fifth embodiment of aspects of the present invention; 
         FIG. 9  is a schematic cross section of a sixth embodiment of aspects of the present invention; 
         FIG. 10  is a schematic cross section of a seventh embodiment of aspects of the present invention; and, 
         FIG. 11  is a schematic illustration of an eighth embodiment of aspects of the present invention. 
     
    
    
     DETAILED DESCRIPTION OF THE EMBODIMENTS 
     With reference to  FIG. 1 , a ducted fan gas turbine engine generally indicated at  210  has a principal and rotational axis XX. The engine  210  comprises, in axial flow series, an air intake  211 , a propulsive fan  212 , an intermediate pressure compressor  213 , a high-pressure compressor  214 , combustion equipment  215 , a high-pressure turbine  216 , and intermediate pressure turbine  217 , a low-pressure turbine  218  and a core engine exhaust nozzle  219 . A nacelle  220  generally surrounds the engine  210  and defines the intake  211 , a bypass duct  222  and a bypass exhaust nozzle  223 . 
     The gas turbine engine  210  works in a conventional manner so that air entering the intake  211  is accelerated by the fan  212  to produce two air flows: a first air flow A into the intermediate pressure compressor  213  and a second air flow B which passes through a bypass duct  222  to provide propulsive thrust. The intermediate pressure compressor  213  compresses the air flow directed into it before delivering that air to the high pressure compressor  214  where further compression takes place. 
     The compressed air exhausted from the high-pressure compressor  214  is directed into the combustion equipment  215  where it is mixed with fuel and the mixture combusted. The resultant hot combustion products then expand through, and thereby drive the high, intermediate and low-pressure turbines  216 ,  217 ,  218  before being exhausted through the nozzle  219  to provide additional propulsive thrust. The high, intermediate and low-pressure turbines  216 ,  217 ,  218  respectively drive the high and intermediate pressure compressors  214 ,  213  and the fan  212  by suitable interconnecting shafts. 
     The compressors and turbines each comprise an annular array of radially extending blades mounted on a rotor disc. Each array of blades may have an annular array of vanes either upstream and/or downstream with respect to the main working fluid passing through the engine. Particularly, the turbine blades and vanes require cooling and the present invention relates to a new cooling arrangement within such a blades and vanes. The present invention may also be applied to compressor blades and vanes. 
     It is known that carefully positioned radially inclined turbulators or ribs in the form of undulations in opposed parts of a passage through which a fluid flows such as a coolant flow passes can generate a rotating vortex as shown in  FIG. 2 . This rotating vortex has a substantial lateral aspect, that is to say rotating laterally to the general longitudinal direction of flow perpendicular to and extending out from the page upon which  FIG. 2  is depicted. By changing the undulations, that is to say rib orientation it is also known that this can generate potentially dual vortices or secondary flows although not of a strong nature. To be effective to improve impingement cooling effectiveness greater flow force is required. As can be seen in  FIG. 2  a component such as a hollow blade  1  has a passage  2  in which opposed parts  3 ,  4  include undulations to generate a rotating or lateral vortex  5  which rotates generally adjacent walls  6  of the passage  2 . The path of the vortex  5  is shown by arrowheads  7 . 
     Fluid flow, that is to say coolant flow from the passage  5  passes through impingement orifices or apertures  8  to project the flow towards a leading passage  9 . The leading passage  9  cools a leading edge of the blade  1  and furthermore includes film orifices  10  which create a coolant film upon the surface of the blade  1  about the lead edge such that in addition to the cooling effect H− the excessive high material temperatures Tm+ are separated from the component  1  through the coolant film generated through the orifices  10 . 
     Although provision of the vortex  7  enhances turbulence and projection flow through the impingement orifices  8  it will be understood that this is not ideal. Directionality as well as further turbulence within the effective feed passage  2  would improve overall performance. By aspects of the present invention a number of vortices are created within the feed passages in accordance with aspects of the present invention. 
     By shaping walls between the undulations or ribs powerful vortices can be generated.  FIG. 3  provides an illustration in which a component in the form of a hollow blade  21  includes a passage  22  having opposed ribs or undulations  23 ,  24 . In such circumstances double vortices  25  are created through a shaped portion  26  in the walls of the passage  22  between the undulations  23 ,  24 . The shaped portion  26  is generally angular in order to provide a division within the passage  22  between the vortices  25   a ,  25   b  to reducing cross flow. 
     It will be understood advantages with regard to providing double vortices  25  in the passage  22  create benefits with regard to:
     a) Increasing the velocity of impingement by jets in the direction of dotted line  11  projected through impingement apertures  28 . Increasing the velocity of the jets  11  will increase the dynamic head at the inlet to the impingement hole. Thus an increase in internal heat transfer in the leading edge passage H+ will occur with a reduced metal temperature at the leading edge Tm−.   b) Increasing the total pressure in a lead passage  29  will also allow the feed flow pressure through the passage  22  to be lowered without reducing the edge film pressure margins through the film apertures  20 . In such circumstances film cooling is more optimal and there is a reduction in leakages from the blade cooling system.   

     As the shaping of the shaped portion  26  is constant it will be appreciated that problems with respect to variability during an operational life for a component will not occur and the shaped portion  26  can be created upon forming the blade  21 .  FIG. 3  provides a schematic cross section of a first embodiment of aspects of the present invention but it will appreciated that other embodiments and variations may be created as described below with respect to other  FIGS. 4 to 11 . Variations can also be achieved through variations in the undulations  23 ,  24 , the shaped portion  26  and the size and orientation of the impingement apertures  28  projecting the flows  11  towards the opposed parts of the leading passage  29 . 
       FIG. 4  provides a further illustration of the embodiment depicted in  FIG. 3  with the circulation arrows etc removed to provide greater detail. It will also be noted that the shaped portion  26  includes further undulations  33 ,  34  to further enhance creation of vortices within the passage  22  in terms of strength and definition. These vortices as indicated before will have a significant lateral aspect in comparison with the flow direction which will generally be perpendicular to the page within which  FIG. 4  is depicted and so along the passage  22 . In such circumstances as described previously more powerful vortices will be created which will be projected towards the impingement apertures.  28  into the leading passage  29  and therefore generate films through film apertures  22  and impingement cooling by engaging opposed parts to a wall portion within which the impingement apertures  28  are created. It will be understood that provision of undulations  33 ,  34  in addition to undulations  23 ,  24  within the confines of the passage  22  may add to manufacturing complexity in comparison with smooth surfaces as depicted in  FIG. 3  but will create as indicated stronger vortices and therefore potentially better cooling effects within a hollow blade component  21 . 
       FIG. 5  provides a schematic cross section of a leading part of a hollow component  41  in which a second embodiment of aspects of the present invention is depicted. As previously a passage  42  includes opposed undulations  43 ,  44  to generate a lateral aspect in a fluid flow, that is to say coolant flow through the passage  42 . The coolant flow will pass longitudinally along the passage  42  and the lateral aspect due to the opposed undulations will be enhanced by a shaped portion  46 . The shaped portion  46  is curved in comparison with the straight angular depictions as shown in  FIG. 3  and  FIG. 4 . Such curvature may enhance vortex generation. Furthermore as depicted by broken lines  143 ,  144  further undulations or ribs may be created in the shaped portion  46  to enhance vortex creation. As previously an impingement wall portion  148  includes impingement orifices or apertures  48 . The impingement orifices  48  project coolant flow generated in the vortices in the passage  42  into and within a leading passage  49 . The leading passage  49  includes film apertures  40  and generally as with previous embodiments includes its own ribs or apertures  149   a ,  149   b  to stimulate turbulence within the leading passage  49  for improved flow turbulence and therefore heat transfer. 
     As illustrated above with regard to  FIG. 3  generally the vortices  25   a ,  25   b  will rotate respectively in substantive isolation in separate parts of the passage  22 . Furthermore the direction of rotation with regard to the respective vortices  25   a ,  25   b  will be centred within their respective parts of the passage  22  to create side by side portions of the fluid flows in the vortices  25 . As illustrated in  FIG. 6  and a third embodiment of aspects of the present invention such an approach allows provision of a single impingement orifice  58  in an impingement wall  158  in a hollow blade component  51 . Thus as previously a passage  52  includes undulations or ribs  53 ,  54  to create a lateral aspect to the fluid flow which has a rotating vortex in accordance with aspects of the present invention and by a shaped portion  56  in the wall of the aperture  52  a number of vortices are generated. The shaped portion  56  as described previously will generate respective vortices which will have side by side components depicted by arrowheads  57  with components  57   a ,  57   b  from each vortex. These components  57   a ,  57   b  will be positioned such that they pass through the impingement orifice  58  into the leading passage  59  for cooling effects as described previously. A single impingement orifice  58  may have advantages with regard to creating a greater flow rate for impingement cooling and pressurisation within the passage  59  and may also facilitate easier fabrication and retain structural strength particularly with a narrow leading edge in the hollow blade component  51 . 
     Although described previously generally with regard to the leading edge of a hollow blade it will also be understood that aspects of the present invention may be utilised with respect to trailing edges of such blades. In such circumstances as depicted in  FIG. 7 , aspects of the present invention comprises a hollow blade component  61  in which a passage  62  acts as a feed passage for coolant fluid flow. The passage  62  includes ribs or undulations  63 ,  64  to generate the lateral vortex flow as described previously and a shaped portion  66  to facilitate vortex creation in respective parts of the passage  62 . The vortices (not shown) will then generate enhanced coolant effects as well as greater impingement flow through an impingement orifice  68  in an impingement orifice wall  168  whereby coolant flow into the trailing edge  69  is enhanced again to improve heat transfer and cooling effects within that passage  69 . In such circumstances it will be understood that aspects of the present invention can be utilised with regard to a trailing edge of a component  61  as well as a leading edge as described previously. 
       FIG. 8  provides a schematic cross section of a leading edge of a hollow blade component  71  including a cooling arrangement in accordance with a fifth aspect of the present invention. Thus, as previously the hollow blade component  71  includes a passage  72  with opposed undulations or ribs  73 ,  74 . In such circumstances again with a fluid flow along the passage  72  lateral flow is stimulated by the undulations  73 ,  74  in order to generate vortices in respective sides of the passage  72 . These vortices enhance flow through impingement apertures  78  in an impingement wall  178  which lead to a leading passage  179  for impingement cooling as well as film development through film apertures  70 . In the fifth embodiment depicted in  FIG. 8  a shaped portion  76  includes shaping towards the front, that is to say the passage  72  as well as the rear for an internal wall which will enhance fatigue life with respect to the shaped portion  76  and therefore generally longevity with regard to operational service life. 
       FIG. 9  provides a sixth embodiment of aspects of the present invention in which only a single passage is employed. In such circumstances a hollow blade component  81  includes a passage  82  in which opposed undulations or ribs  83 ,  84  are provided to generate a lateral vortex flow which through a shaped portion  86  substantially between the undulations  83 ,  84  is further stimulated into providing vortices for enhanced directional flow towards film orifices  80 . In such circumstances the strong vortices created by the shaped portion  86  will have a direct effect upon the film developed through the film orifices  80 . Undulations/ribs could also be added to shaped portion  86  to further enhance the strength of the vortices. 
       FIG. 10  provides a schematic cross section of a seventh embodiment of aspects of the present invention in which again a hollow blade component  91  includes a passage  92  within which opposed undulations or ribs  93 ,  94  act upon a flow through the passage  92  to create lateral vortex aspects which are enhanced by a shaped portion  96  to define the vortices as described previously. In the seventh embodiment depicted in  FIG. 10  a rear surface of the impingement wall  198  is also shaped to enhance and facilitate vortex definition. In such circumstances impingement orifices  98  in the wall portion  198  direct impingement flows towards a leading passage  99 . Impingement flows have generally relatively greater force and pressurisation within the leading passage  98  for enhanced heat transfer and cooling effects within the hollow blade component  91 . As described previously coolant flow from the leading passage  99  passes through film apertures  90  to develop film cooling effects about the leading edge of the component  91 . By providing shaping to both the shaped portion  96  and a rear surface of the wall portion  198  a combination is created with enhanced vortex definition effects from the rotational vortex generated by the opposed undulations or ribs  93 ,  94 . 
     It will be appreciated that shaping to both the passage wall portions to either side of the proposed undulations or ribs in a passage in accordance with aspects of the present invention has greater enhanced effects with regard to vortex creation. In such circumstances, and as depicted in an eighth embodiment of aspects of the present invention shown in  FIG. 11 , a hollow blade component  101  with a passage  102  has a shaped portion  106  and opposed undulations  103 ,  104 . The shaped portion  106  has two raised sections which are opposed by reciprocal parts of the rear surface of the impingement wall portion  208 . In such circumstances with double shaping as illustrated three vortices  105   a ,  105   b ,  105   c  which by their rotational direction engage mostly respective impingement orifices  108  leading to passage  109 . The greater coolant flow pressure in the passage  109  enhances cooling effects and also film development through film orifices  100 . The increased number of holes ( 108 ) also increases the cooling effectiveness due to the greater surface area covered by the jets. 
     It will be appreciated from the above that aspects of the present invention utilise and enhance through shaped portions the rotational vortex or lateral vortex flow aspect generated by opposed undulations or ribs in a general feed passage for a hollow blade component. By shaping portions of the passage vortices of a stronger and tighter aspect are generated which can then be utilised to present stronger flows through impingement orifices to a leading passage or directly to film orifices for enhanced cooling effects in comparison with the coolant flow rate utilised. Such relative enhancement of cooling efficiency will provide significant overall benefits with regard to engine operational performance in that greater cooling effect is achieved allowing increased metal reduction temperatures proportionately or higher operating temperatures with less coolant flow. 
     Aspects of the present invention may be utilised with regard to cooled turbine blades or nozzle guide vanes in a gas turbine engine. These engines may be used in civil, military, marine or industrial applications but by allowing the engine to operate at higher temperatures proportionately to the coolant flow overall operational efficiency is achieved whilst maintaining operational life. As indicated above modifications and alterations to aspects of the present invention may be achieved by a person skilled in the technology. As described the undulations or turbulators in the form of ribs in addition to being in opposed parts of the passage itself may be added to the shaped portions, that is to say the angular walls to increase or optimise the vortex effects and so increase impingement and other cooling effects. 
     The shaped portions may be angular and have flat planar surfaces for sharper definition of sides to the passage or alternatively as illustrated above may be smoothly shaped to increase and again optimise vortex effects. Similarly, undulations or ribs can be presented and formed in the shaped surfaces where required. 
     The number of impingement holes, their position and angles may be altered to achieve higher or lower flow rates in portions and sections opposing the impingement holes in the leading passage for relative local cooling effects thereat. 
     By combining radial and/or tangentially inclined impingement holes the benefits of enhanced vortex control through the shaped portions can be further optimised through flow pickup and direction. 
     Although of particular benefit with regard to leading edges where high temperature problems persist it will also be understood that cooling arrangements in accordance with aspects of the present invention may be utilised in other regions of a blade or aerofoil such as a trailing edge. 
     The rear surface of the shaped portion may be angled or shaped to form a diamond or thicker aspect to increase fatigue life for a blade. It will be understood that such an approach may allow aspects of the present invention to be utilised in situations where there is relatively high stresses and therefore predicted shorter operational life than would be acceptable particularly with the impingement holes as described above. 
     By utilising angled walls in a radial leading passage wall including the impingement orifices it is possible to further increase cooling effectiveness and heat transfer by extending the impingement orifice length and therefore jetting effects with regard to angling as well as enhanced vortex generation within the passage in accordance with aspects of the present invention. 
     By appropriate multiple shaping and angling of the shaped surfaces in accordance with aspects of the present invention multiple vortexes can be created. These vortexes may be substantially all of the same size or have different sizes and vortex strengths if possible through the shaped portions nevertheless, consideration of potential unbalance within the passage may create instability. Such instability may be detrimental to impingement coolant flow force through the impingement holes in accordance with aspects of the present invention. 
     As indicated above generally undulations in accordance with aspects of the present invention comprise ribs formed within the passages. Alternatively, there may be surface treatments to alter the flow friction effects and therefore actions which may provide similar flow control effects to ribs or undulations as described above. 
     In summary of the present invention, an aerofoil of a vane or blade of a gas turbine engine comprises an internal passage through which a cooling fluid passes. The passage is partly formed by first and second opposing walls  27 ,  26  and as shown in  FIGS. 3-11  further defined by the external walls of the aerofoil. The first wall  27  comprises at least one aperture  28  and the second wall  26  comprises angled wall portions  26   a ,  26   b  forming a tip region  26   t  adjacent the first wall. The tip is closest the first wall and the wall portions are divergent away from the first wall. The passage also comprises ribs  23 ,  24 ,  33 ,  34  which together with the wall portions  26   a ,  26   b  create at least two vortices  25   a, b  in the coolant fluid. These vortices rotate such that their direction of rotation forces additional coolant through the apertures to increase the dynamic head of cooling fluid through the aperture. This increases the amount of coolant through the apertures and can improve the impingement cooling of an external wall of the aerofoil. 
     It should be appreciated that the vortices (e.g.  25   a ,  25   b ) extend across their respective portions (e.g.  35   a ,  35   b ) of the passageway  22 . These vortices are rotations of the bulk coolant flow through the passage portions rather than any smaller and local vortices. 
     In  FIG. 3 , the first wall comprises two apertures  28 , although these can be part of a radially extending array of apertures, and they are arranged either side of the tip region  26   t . Although, with two counter rotating vortices which can coalesce to pass through just one aperture (or radial array of apertures), in this preferred embodiment each of the vortices feeds coolant into each of which array of apertures. 
     The ribs are angled relative to a radial line from the engine&#39;s rotational axis and as the coolant passes along the passage it is caused, by the angled ribs, to rotate and form the vortices. The vortices are contained within each portion of the passage by the angled walls  26   a  and  26   b  so that stronger vortices are formed. The ribs are preferably formed on the external aerofoil walls  21 , however, the ribs a can be arranged on any one or more of the walls depending on preferred vortex strength and aerofoil configuration, such as use in a vane or blade and also the position within the aerofoil and its coolant flow quantities. 
     The dynamic head of the coolant flow is increased to provide improved impingement cooling via the apertures. This is particularly, desirable for cooling the inner surface of an external wall subject to the very hot working gases passing through a turbine for example. However, in other applications it, may be desirable to increase the dynamic head through apertures to increase the effectiveness of a cooling film over the aerofoil&#39;s external surfaces and in this case the first wall  27  is an external wall  81 . This is shown in  FIG. 9 . 
     Further detailed improvement can be seen in  FIGS. 10 and 11 . In  FIG. 11 , the second wall  106  comprises more than one pair of angled wall portions  106   a, b, c, d  forming a number of tip regions  106   t  positioned adjacent the first wall  107 . This arrangement creates three or more vortices  105   a, b, c  in the coolant fluid which are themselves adjacent and feeding corresponding apertures  108  in the first wall  107  to increase the dynamic head of cooling fluid through the aperture. 
     In  FIGS. 10 and 11 , the first wall  107 ,  97  comprises one or more pairs of angled wall portions  97   a, b ,  107   a, b, c, d  which form a number of tip regions  97   t ,  106   t  positioned near to the adjacent the second wall  26 . The opposing tip regions  97   t ,  106   t  of the first wall  27  and tip regions  26   t ,  97   t ,  106   t  of the second wall  26  are adjacent one another and help retain and increase the strength of the vortices. 
       FIG. 5  shows the wall portions  46   a ,  46   b  are concave, but they could be straight or another arcuate form to improve the strength of the vortices.