Passive clearance system for turbine blades

Spent cooling air from the internal passage(s) in a turbine blade of a gas turbine engine are judiciously located to inject air at the pressure side of the blade in proximity to or in the gap between the tip of the blade and outer air seal to reduce leakage in the gap and improve engine performance. In one embodiment, a projection at the tip of the blade on the pressure side is utilized in combination with an angled discharge passageway and in another embodiment a pair of discharge angled passages, one interconnecting an internal passage on the pressure side and the other connecting an internal passage on the suction side of the blade is utilized. A projection at the tip adjacent the pressure side may also be incorporated in the second embodiment.

TECHNICAL FIELD 
This invention relates to air cooled turbine blades for gas turbine engines 
and more particularly to means for utilizing the cooling air in 
combination with the air adjacent the surface of the blade for clearance 
control. 
BACKGROUND ART 
As is well known, leakage flow from the pressure side to the suction side 
of a turbine blade across the tip (the gap between the blade tip and the 
outer air seal) results in a performance loss of aircraft gas turbine 
engines. There has been over the years a continual effort to maintain a 
close clearance between the tips of an axial flow turbine blades and the 
outer air seal surrounding these tips for the entire operating envelope of 
the engine. As one skilled in the art appreciates, because the rotor has a 
greater mass than the engine casing, the rotor will expand and contract in 
response to temperature changes slower than the casing. The engine is 
initially designed so that the tips of the blades will not rub against the 
outer air seal for both transient and steady-state conditions. Hence, the 
gap must be sufficiently large to accommodate certain transient conditions 
and yet be small when the engine is operating at a steady-state condition. 
This presents problems since the gap is designed to obviate rubbing to 
accommodate the transient conditions. When the engine returns to the 
steady-state condition the gap is generally larger than desired unless 
means are taken to adjust for this problem. This problem is acerbated when 
an engine, particularly powering military fighter aircraft, is put through 
extreme transient conditions such as throttle chops, rapid re-accels and 
the like which require the engine case and rotor components to respond 
more rapidly than would otherwise be the case in a commercial airline. 
Presently, there are two fundamental ways in which this gap is controlled, 
one by an active control system and the other by a passive control system. 
Essentially, an active control system, sometimes referred to as active 
clearance control, typically relies on some external heat or cooling 
source and the actuation of an external control system that serves to 
conduct the heating or cooling from the source to the component parts in 
proximity to the blade so as to change their temperature in order to 
effectuate contractions or expansion of the involved components and hence 
change dimension of the gap. A passive system, on the other hand, relies 
on the surrounding environment to effectuate the gap closure. Examples of 
an active clearance control can be had by referring to U.S. Pat. No. 
4,069,662, granted to Redinger et al on Jan. 24, 1978 and assigned to the 
assignee common to this patent application. Examples of a passive 
clearance control system can be had by referring to U.S. Pat. Nos. 
3,575,523, granted to F. J. Gross on Apr. 20, 1971; 4,534,701 granted to 
G. Wisser on Aug. 13, 1985 and 4,863,348 granted to W. P. Weinhold on Sep. 
5, 1989. 
One method of reducing the leakage of air across the tip of the turbine is 
to discreetly inject the discharge air from the turbine blades internal 
cooling passages at judicious locations at the tip of the blade adjacent 
the pressure side of the blade. This serves to create a buffer zone and 
forms a curtain of air to effectively minimize the leakage occurring 
across the tip of the blade from the pressure side to the suction side. 
DISCLOSURE OF THE INVENTION 
An object of this invention is to provide for an axial flow turbine blade 
of a gas turbine engine improved passive means for reducing leakage of 
engine working medium adjacent the tip of the turbine blade. 
A feature of this invention is the location of an axial projection located 
on the pressure side adjacent the tip of the blade that together with the 
discreet discharge of cooling air from the blade and the radial air flow 
adjacent the outer surface of the blade on the pressure side provides a 
"curtain" of air adjacent the tip of the blade to minimize leakage of the 
engine working medium at the tip of the blade. 
A still further feature of this invention is to route the air internal of 
the blade adjacent the suction side of the blade to judiciously discharge 
air adjacent the tip of the blade at a discreet low angle. 
A still further object of this invention is to provide a passive tip 
leakage reducing system at the trailing edge utilizing straight holes 
rather than curved holes that have been used heretofore. 
The foregoing and other features and advantages of the present invention 
will become more apparent from the following description and accompanying 
drawings.

BEST MODE FOR CARRYING OUT THE INVENTION 
While only two specific embodiments of turbine blades are disclosed herein, 
it will be understood that many of the internally air cooled turbine 
blades can utilize this invention. Additionally, for the sake of 
convenience and simplicity only a portion of the turbine blade is 
disclosed and for more details of a suitable blade reference is hereby 
made to the F100 family of gas turbine engines manufactured by Pratt & 
Whitney division of United Technologies Corporation, the assignee of this 
patent application. Suffice it to say, and referring to FIG. 1 the turbine 
blade generally illustrated by reference numeral 10 is one of a plurality 
of blades suitably supported in a disk and rotably mounted on the engine 
shaft for powering the engine's compressor (not shown). The engine's fluid 
working medium (gas generated by the burner) serves to impinge on the 
airfoil of the blade, which in turn, extracts a portion of its energy for 
driving the compressor while the remaining energy is utilized to develop 
engine thrust. 
As best seen in FIG. 2, cooling fluid internally of the blade discharges 
from a plurality of holes 14 (one being shown) into the engine's fluid 
working medium from an internal cavity 12. As noted, the secondary flow, 
illustrated by arrow B, which is the engine's working medium adjacent the 
blades surface 16 on the pressure side 18 of the blade, travels in a 
radial direction relative to the engine's center line and combines with 
the flow discharging from orifice 14. 
In accordance with this invention, a small projection 20 extending axially 
in the aft direction relative to the main stream of the engine's working 
medium is located at the tip 22 of blade 10. The holes 14 are drilled 
parallel to the angular wall 24 and are substantially equal to an angle 
that lies between and including 35.degree. to 60.degree. relative to the 
projected center line of the engine as viewed in FIG. 2 (designated by 
reference letter X) or relative to the surface of tip 22. The combined 
flow, i.e., the radial secondary flow and the discharge flow, serve to 
form a curtain of air adjacent the tip 22 and block the leakage flow 
illustrated by arrow A. The leakage flow that flows in the gap 28 formed 
between the tip 22 and outer face 30 of the outer air seal 32 is in the 
direction from the pressure side 18 to the suction side 34. This curtain 
of air adjacent the inlet end of the gap, reduces the amount of leakage 
flow that would otherwise occur. The reduction in leakage of the high 
energy fluid working medium causing the otherwise leakage flow to pass 
through the working surface of the blade enhances turbine performance and, 
hence, engine performance. 
The second embodiment exemplified in FIGS. 3 and 4 is generally similar to 
the embodiment disclosed in FIGS. 1 and 2 save for the extended projection 
38 corresponding to projection 20 of FIGS. 1 and 2 and the inclusion of 
internal cavity 48 adjacent the suction side 52. Referring more 
specifically to FIGS. 3 and 4, the blade generally illustrated by 
reference numeral 44 is bounded by the trailing edge 40, leading edge 42, 
the pressure side 50 and the suction side 52. For the purpose of this 
description, the portion of the blade extending between the dimensions 
denoted by the arrow D is the trailing edge region, the arrow E is the 
mid-portion region and the remaining portion is the leading edge region. 
In accordance with this invention, the cooling air from the cavity 48 which 
is utilized for blade internal cooling is routed to the tip 56 of blade 44 
through drilled holes or passageways 58. The passageway 58 is oriented 
such that the flow of air is injected at an angle that is relatively low, 
say between 35.degree. and 60.degree. with respect to the projected 
engine's center line (designated by reference letter Z) or with respect to 
the surface of tip 56. The injection of the air at this low angle serves 
to provide at the pressure side 50 a curtain of air adjacent gap 60 formed 
between tip 56 and the inner surface 62 of the outer air seal 64. 
In certain embodiments, it would not be practical to provide such a curtain 
in the tip region because of the inherent design of the cooling passages 
in the blade. 
As is apparent for the description above, the internal passages in the 
blade 44 include both the cavity 48 adjacent the suction side 52 and 
cavity 46 adjacent the pressure side 50. Both cavities extend radially in 
blade 44 and carry cooling air and are in proximity to the tip portion of 
blade 44. In this embodiment, air is directed to form a curtain at the tip 
56 adjacent pressure side 50 by injecting air at a low angle from cavity 
46 through drilled hole or passageway 54 and from cavity 48 through 
drilled hole or passageway 58 the drilled hole 58 is at a shallow angle 
similar to the angle of the slope of the projection 38 which angle is 
substantially equal to and including between 35.degree. and 60.degree. 
relative to the projected engine's center line or relative to the surface 
of tip 56. This assures that there will be a continuous sheet of high 
velocity air adjacent gap 60 to oppose leakage flow therein. A plurality 
of passageways 54 and 58 (only one of each being shown) preferably should 
be staggered along the length of the pressure side of the blade. 
As is apparent from the foregoing, the cooling air which has been used to 
cool the blade 44, is utilized further for minimizing leakage occurring in 
gap 60. In one instance, the spent cooling air from the pressure cavity 46 
(one being shown although some blades may have a plurality of such 
cavities) is combined with the natural radial secondary flow along the 
pressure side face and combined with the spent cooling air from the 
suction side cavity 46 (only one being shown although some blades may have 
a plurality of such cavities) and in the other instance a projection at 
the tip 56 is utilized depending on whether or not such a projection would 
be practical on the leading edge region of the blade. 
In this embodiment, the injection of spent cooling air adjacent gap 60 is 
at a relatively low angle and is available from two separate cavities, 
located on both the pressure and suction sides of the blade for reducing 
tip leakage. 
Although this 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.