Patent Abstract:
Where gas turbine engine structure eg combustion equipment, is to be air impingement cooled, the surface which receives the air jets is so shaped as to produce boundary layer separation zones  34, 38  and  44  in the cooling air, as it spreads across the surface. Mixing of the boundary layer with the remainder of the air flow results, followed by the re-establishment of the boundary layer. The new boundary layer is cooler than the original layer and so provides more effective cooling.

Full Description:
This is a Continuation-in-Part of National Appln. No. 09/748,861 filed Dec. 28, 2000 abandoned. 
    
    
     FIELD OF THE INVENTION 
     The present invention relates to a system for cooling components which in use, experience high temperatures. The invention has particular efficacy in the gas turbine field, and may be incorporated in gas turbine engines of the kinds used to power aircraft or ships, or to pump oil overland. 
     BACKGROUND OF THE INVENTION 
     Air impingement cooling of gas turbine engine combustion equipment and other structures therein, is well known. However, known systems, wherein cooling air flowing over the surface of one member, passes through holes and crosses a gap, to impinge on a surface of an adjacent hot member, fail to achieve their full cooling potential. This is because the jet of air, on striking the surface of the hot member, spreads over the surface, effectively in a layer of constant thickness. It follows, that the outer portion of the layer never touches the hot member, and consequently, cannot make an efficient contribution to the cooling effect of the air flow. 
     A further drawback to known impingement cooling systems, is that, having impinged on the hot surface, and spread through 360° over the hot surface, the respective air flows collide with each other, and form a turbulent mix with poor heat transfer performance, and which sometimes displaces incoming air jets. Hot spots are thus formed. 
     SUMMARY OF THE INVENTION 
     The present invention seeks to provide an improved air impingement cooling system. 
     According to the present invention, an air impingement cooling system comprises superimposed, spaced apart members, one perforated, the other having a surface portion directly under each respective perforation, each said surface portion being of fluctuating shape, so as to cause air received thereby via respective perforations, and deflected laterally there across, to flow over said fluctuations, said fluctuating shape being such that the boundary layer of said air flow over said surface portion is caused to separate from said surface portion in the region of said fluctuations and subsequently reform downstream of said separation. 
    
    
     BRIEF DESCRIPTION OF THE DRAWINGS 
     The invention will now be described, by way of example, and with reference to the accompanying drawings in which: 
     FIG. 1 is a diagrammatic view of a gas turbine engine having combustion equipment which incorporates the present invention. 
     FIGS. 2 to  6  are examples of alternative configurations of the present invention. 
     FIG. 7 is a view in the direction of arrow  7  in FIG.  6 . 
    
    
     DETAILED DESCRIPTION OF THE INVENTION 
     Referring to FIG. 1. A gas turbine engine  10  has a compressor  12 , combustion equipment  14 , a turbine section  16 , and an exhaust nozzle  18 , all arranged in flow series in known manner. The operation of the gas turbine engine  10  is well known and will not therefore be described herein. 
     The combustion equipment comprises flame tubes  20 , surrounded by a casing  22 , which is spaced therefrom. The space is numbered  24 . Casing  22  is itself spaced from an outer engine casing  26 , that space being numbered  28 . 
     Space  28  is connected to receive a flow of air from compressor  12 , which air flows over the outer surface of casing  22 , some air thus by-passing the flame tubes  20 , the remainder passing through a large number of holes  19  in casing  20  (FIGS. 2-6) to impinge on the outer surface of respective flame tubes  20 , so as to cool them. The air is in the form of individual jets, numbered  30 . (FIGS.  2 - 6 ). 
     Referring to FIG. 2, in this example, when an air jet  30  strikes the outer surface portion of flame tube  20  which is directly under it, the air spreads laterally of the jet, over 360° across that surface portion, until it meets a barrier defined by wall  32 , which totally bounds the surface portion struck by and expanded over by the air jet, up to the limit where, without the presence of the wall  32 , the spreading flow would collide with those flows spreading from immediately adjacent jets. Thus, the wall  32  completely surrounds the surface portion as is the case in the FIGS. 3,  4  and  5  examples. Additionally, each surface portion bounded by a wall  32  is impinged by an air jet  30  from a single hole  19  the axis  21  of which intercepts the surface portion substantially at the center of each surface portion. Also, as shown in FIG. 2, the following dimensional relationships may be employed where d is the diameter of the hole  19  and h is distance from the casing  22  to the surface portion  34  bounded by the wall  32  which may slope at an angle α from the surface portion and the distance from the point of interception of the axis  21  of the hole  19  to the boundary wall  32  is L: L≧d; α≧30°; the height of the wall  32  should be≦0.3h. 
     On striking the wall  32 , the boundary layer of the cooling air flow, that is, the portion of the flow immediately adjacent the surface portion, separates from the surface portion in the region  34 . This causes mixing of the boundary layer and the remainder of the cooling air flow, before the boundary layer reforms and attaches itself to the wall. However the reformed boundary layer is cooler than the previous boundary layer due to this mixing and so provides more effective cooling of the wall  32 . 
     On perusal of FIGS. 2 to  5 , it will be clear to the expert in the field, that the wall  32  also provides parts of boundaries for those jets immediately surrounding the jet  30 , an example being depicted in FIG. 7, to which reference is made later in this specification. 
     Referring to FIG.  3 . in which like parts have like numbers. In this example, the centre of the portion bounded by wall  32  is provided with a cone  36 , the apex of which faces into the jet  30 . Such a shape defines a fluctuation in surface shape at its junction with the flame tube  20  outer surface. This fluctuation causes separation of the boundary layer flow in the region  38 . The separated boundary layer, which at this position is hotter than the remainder of the cooling air flow, mixes with, and is thereby cooled, by the remainder of the cooling air flow. A new, cooler and thinner boundary layer then forms which proceeds to flow towards the wall  32 , in turn providing more effective cooling of the outer surface of the flame tube  20 . 
     Referring to FIG.  4 . In this example, separation of the boundary layer of the cooling air flow is provided in the region  38  by the provision of a rising slope  42  in the surface portion. The separated boundary layer then mixes, and is therefore cooled, by the remainder of the cooling air flow before a new, cooler, boundary layer is formed which flows towards the wall  32 . 
     Referring to FIG.  5 . This example combines the cone  36  of FIG. 3 with the rising slope  42  of FIG. 4, and produces, in the one arrangement, boundary layer separation which occurs in the regions  34 ,  38  and  44 , thereby providing more efficient cooling. 
     Referring to FIG.  6 . This example utilises the rising slope  42  of FIG. 4, but not the boundary wall  32  thereof. Instead, the rising slope  42  of FIG. 6 meets rising slopes eg  42   a  and  42   b  of adjacent surface portions, which features are more clearly seen in FIG.  7 . The advantages accrued by the arrangement depicted in FIG. 6 are reduction in weight, and at least a reduction in turbulence, when opposing, spreading air flows meet, by virtue of the flows already having a small directional component, which will serve to generate a resultant direction of flow of the collided air flows, in parallel with the jets. 
     Referring now to FIG.  7 . When opposing, spreading air flows collide, they tend to form a barrier which approximates a straight line. Thus, ridges  46  represent that line, one such ridge  46  lying between the heads of respective groups of arrows  48  and  50 , which in turn, represent colliding air flows. From this, it will be appreciated that each impingement surface is bounded by a plurality of straight lines which, in the present example, define a pentagon. 
     However, in practice of the present invention, the actual number of straight lines and therefore, the shape defined, will be dependant on the number of perforations  19  in casing  20  (not shown in FIG. 7) and the pattern in which they are drilled. 
     Boundaries of circular shape (not shown) may be provided, but the resulting interstices of solid metal would add weight. If they were to be machined out, cut-outs would have to be made in the boundary edges, so as to allow spreading cooling air to flow into the resulting pockets. 
     The cone  36  in both FIG.  3  and FIG. 5 may be of circular form in cross section. Alternatively, it could be multi-faceted e.g. pyramid-like.

Technology Classification (CPC): 5