Technical Field

This invention relates to coolable rotor blade structures of the type used in high temperature rotary machines, and more specifically, to structure for reducing stresses in such airfoils.

Background Art

An axial flow rotary machine, such as a gas turbine engine for an aircraft, includes a compression section, a combustion section and a turbine section. A flow path for hot working medium gases extends axially through the engine. The flow path for hot gases is generally annular in shape.

As working medium gases are flowed along the flow path, the gases are compressed in the compression section causing the temperature and pressure of the gases to rise. The hot, pressurized gases are burned with fuel in the combustion section to add energy to the gases. These gases are expanded through the turbine section to produce useful work and thrust.

The engine has a rotor assembly in the turbine section which is adapted by a rotor disk and blades extending outwardly therefrom to receive energy from the hot working medium gases. The rotor assembly extends to the compression section. The rotor assembly has compressor blades extending outwardly across the working medium flow path. The compressor blades rotate with the rotor assembly and drive the incoming working medium gases rearwardly, compressing the gases and imparting a swirl velocity to the gases. The high-energy working medium gases in the turbine section are expanded through the turbine blades to drive the rotor assembly about its axis of rotation. Rotational forces acting on the rotor blade as the rotor blade is driven about the axis of rotation cause stresses to which the blade.

Each rotor blade in the turbine section has an airfoil to receive work from the gases to direct the hot working medium gases through the stage of rotor blades. As a result, the airfoils are bathed in hot working medium gases during operation causing

thermal stresses in the airfoils. These thermal stresses in combination with the stresses of rotation unit the structural integrity and fatigue life of the airfoil.

Accordingly, rotor blades are typically cooled to reduce thermal stresses and thereby provide the rotor blade with a satisfactory structural integrity and fatigue life. An example of such a rotor blade is shown in US Patent 4, 474,532 entitled

"Coolable Airfoil For a Rotary Machine", issued to Pazder and assigned to the assignee of this application. Another example of a coolable rotor blade is shown in US Patent 4, 278, 400 issued to Yamarik and Levengood entitled "Coolable Rotor Blade" and assigned to the assignee of this application. Each of these rotor blades is provided with a plurality of cooling air passages on the interior of the blade. Cooling air is flowed through the passages to the rearmost portion of the rotor blade, commonly referred to as the trailing edge, from whence the cooling air is exhausted into the working medium flow path.

The above art notwithstanding, scientists and engineers working under the direction of Applicant's assignee are seeking to develop coolable airfoils for use in high temperature rotary machines which have acceptable level of stresses in critical regions of the airfoil.

Disclosure of Invention

According to the present invention, a coolable rotor blade having an airfoil includes a radially extending portion in which the center of gravity of the radially innermost airfoil section is rearwardly of the center of gravity of the remainder of that portion of the airfoil.

This invention is in part predicated on the recognition that the maximum stress that occurs in the trailing edge region is a tensile stress under operative loads and occurs at the junction between the trailing edge and the root region of the airfoil and, that a portion of the stress may be shifted to the lending edge region.

In accordance with one embodiment of the present invention, the spanwisely extending portion of the airfoil extends from the root region of the airfoil to the tip of the airfoil, the trailing edge region is planar along the pressure side surface and the trailing edge is curved forwardly over the entire portion of the airfoil in a plane parallel to the planar pressure side surface.

A primary feature of the present invention is an airfoil having a radially extending portion. The radially extending portion has a plurality of airfoil sections, the center of gravity of each of which is axially rearward of the center of gravity of the remaining portion of the airfoil sections. In one detailed embodiment, the coolable airfoil has a plurality of discharge slots in the trailing edge region of the airfoil. In another detailed embodiment, the trailing edge region is planar on the pressure side surface and is curved forwardly in the radially outward direction when viewed in a direction perpendicular to the generally planar surface.

A primary advantage of the present invention is the level of stresses in the trailing edge region which results from the compressive stress exerted under operative conditions by the forwardly displaced center of gravity of the radially extending portion of the airfoil. Another advantage is the ease of manufacture which results from the generally planar surface on the pressure side wall of the airfoil in the trailing edge region which enables a planar abrasive surface to uniformly remove material during finishing operations of the airfoil along the radial length of the airfoil. Another advantage is the flexibility in blade design which results from controlling the amount of stress reduction over the trailing edge region of the airfoil by controlling the trailing edge curvature of the airfoil.

Other features and advantages will be apparent from the specification and claims and from the accompanying drawings which illustrate an embodiment of the invention.

Brief Description of Drawings

Fig. 1 is a partial perspective view of a portion of a rotor assembly for an axial flow rotary machine; Fig. 2 is a side elevation view of the rotor blade shown in Fig. 1 with a portion of the rotor blade broken away for clarity along a curved sectional line through the airfoil;

Fig. 3 is a top view taken along the lines 3.3 of Fig. 2 showing an airfoil section and the center of gravity of the radially outward portion of the airfoil shown in phantom;

Fig. 4 is a view taken generally along the line 4 of Fig. 3 showing the planar trailing edge region of the airfoil;

Fig. 5 is a view taken along the line 5 of Fig. 3 in a direction generally perpendicular to the planar trailing edge surface of the airfoil.

Best Mode for Carrying out the Invention

A gas turbine engine embodiment of an axial flow rotary machine is illustrated in the Fig. 1 partial perspective view. A portion of a rotor assembly 10 of the engine is shown. The rotor assembly includes a rotor disk 12 having an axis of rotation R and a plurality of rotor blades 14 extending outwardly from the disk. A flow path 16 for working medium gases extends through the rotor assembly passing between adjacent rotor blades.

Each rotor blade 14 has a root 18, a platform 19 an airfoil 20 and a tip 22. Dimensions of the rotor blade are measured from a reference X-plane, a reference Y- plane and a reference Z-plane. In the installed condition, these reference planes have a particular orientation with respect to the axis of rotation R of the rotor assembly. The reference X-plane extends in the axial direction and contains the axis of rotation R. The reference Y-plane is a radial plane perpendicular to the axis of rotation R. The X and Y planes intersect along a reference line, commonly called the stacking line L... The reference Z-plane is perpendicular to both the X and Y planes. The Z-plane intersects the stacking line and is a tangent plane at an arbitrary radius from the axis of rotation R. The X, Y, and Z planes remain with the blade in its uninstalled condition. As shown in Fig. 2 the root has a first duct 24 and a second duct 26 which are in fluid communication with a source of cooling air. One source of cooling air is the compression section of the engine (not shown). The airfoil, 20 has a leading edge 28 disposed in the leading edge region 32 and a trailing edge 34 disposed in the trailing edge region 36. A suction sidewall 38 and a pressure sidewall 42 (partially broken away in Fig. 2 for clarity and shown in Fig. 3) are joined at the leading edge and the trailing edge. The pressure sidewall is spaced from the suction sidewall to form a cavity 44 therebetween. Internal walls divide the interior of the cavity. An internal tip wall 46 extends chordwisely between the pressure sidewall 42 and the suction sidewall 38 to bound

the cavity 44 in the spanwise direction. An internal wall such as a first rib 48 extends in a spanwise direction. The first rib is spaced from the leading edge 28 to form a first cooling air passage 52 there between. It is spaced spanwisely from the tip wall 46 to leave a turning passage 54 there between. A second rib 56 extends spanwisely from the tip wall into proximity with the root 18 to define an inwardly extending passage 58 in flow communication with the turning passage 54 and a discharge passage 62 in flow communication with the trailing edge region 36 of the airfoil. The trailing edge region of the airfoil has a plurality of pedestals 64 and lands 66 bounding slots 68. The slots 68 extend rearwardly through the trailing edge to place the trailing edge region in flow communication with the exterior of the rotor blade.

The pressure sidewall surface in the trailing edge region is generally planar so that the trailing edge may be ground to its final contour with the use of a planar grinding surface. One example of a generally planar surface is one which deviates ten spanwisely about thirteen toH_ε_fed thousandths of an inch per inch (.0013) at the trailing edge and sixty-six htmd-F∞l thousandths or an inch per inch (.0066) at a chordwise location which is about one hundred and seventy-five thousands (.175 inches) from the trailing edge. As will be realized, this greatly increases th manufacturability of the blade and facilitates the accuracy of the grinding of the fina contour to preserve the aerodynamic efficiency of the rear of the blade.

The airfoil section 72 taken along the line 33 of Fig. 2 shows the pressur sidewall 42 and the suction side wall 38. The sidewalls are tangent to a trailing edg cylendrical surface at the trailing edge.

The airfoil is formed of a plurality of such airfoil sections. Each airfoi section extends generally perpendicular to the radial line. Each airfoil section extend parallel to the remaining airfoil sections and has a center of gravity as represented b the center of gravity CGI. A radially outward airfoil section contains the center o gravity for the remaining portion of radially outwardly airfoil sections between th airfoil section shown and the tip 46 of the airfoil. Thus, the center of gravity of th airfoil section shown in Fig. 3 is rearwardly of the center of gravity of the remainde of the spanwisely extending portion of the airfoil as measured in a plane parallel t the pressure sidewall surface in the trailing edge region.

Fig. 3 is a view taken along the line 3-3 of Fig. 2 with the airfoil 20 partly in section and partly in full with details of the section broken away for clarity.

As can be seen in Fig. 4, which is a view taken generally along the line 4 of Fig. 3, the trailing edge region is generally flat and straight because of the planar surface on the pressure side wall of the rotor blade.

Fig. 5 is a view taken along the line 5 of Fig. 3 in a direction which is perpendicular to the planar surface of the pressure sidewall of the trailing edge region. As shown in Fig. 5, the trailing edge is curved forwardly in the radially outward direction. During operation of the rotor disk-blade assembly shown in Fig. 1, the rotor assembly is driven about is axis of rotation R,. As a result of this rotational movement, centrifugal forces act on the rotor blade causing a tensile stress in the trailing edge region 36 of the rotor blade. The combination of tensile stresses and thermal stresses which results from the working medium gases flowing through the airfoil is a maximum at the joinder of the trailing edge 34 to the region of the airfoil adjacent the platform 19 of the rotor blade.

Because of the rotational moment of forces exerted on the center of gravity of each of airfoil section, each airfoil section, exerts a compressive stress at the trailing edge of the airfoil on each radically inward airfoil section of the airfoil. This reduces the tensile stresses in the airfoil of the trailing edge region in comparison with airfoils which do not have this particular orientation of the center of gravity of the airfoil sections. This is particularly desirable in this construction because the trailing edge of the cooled turbine blade is hotter than the body of the blade and is subject to stress concentrations associated with the trailing edge cooling slots and thus benefits from reduced stress levels relative to the rest of the blade. The leading edge is cooled by the cooling air passage and can better accommodate the tensile stresses associated with this construction.

In the particular embodiment shown, the blade tailing edge is curved enabling stacking of the airfoil sections relative to a radial line to control the amount of stress reduction over the entire span of the blade. Thus, varying the movement of the airfoil sections and the center of gravity for each airfoil section in a plane parallel to the plane of the pressure side surface varies the curvature of the trailing edge and

varies the reduction in the amount of stress. And, the variation in curvature may be changed in different designs to reduce tensile stress in that portion of the trailing edge region which most requires the reduction. Inaddition, blade aerodynamics are not compromised in this particular design. In this design, airfoil sections are shifted by shifting airfoil cross sections that are designed to aerodynamically optimum, not by changing the shape of the cross sections. Trailing edge stress reduction and a straight trailing edge could probably be achieved by changing the shape of the airfoil cross sections, but this would compromise blade aerodynamics.

Although the invention has been shown and described with respect to detailed embodiments thereof, it will be understood by those skilled in the art that various changes in form and detail thereof may be made without departing from the spirit and scope of the claimed invention.