Method and system for the removal of large turbine engines

A detachable support structure for a large turbine aircraft engine having a nacelle which is attached to an auxiliary support frame. The auxiliary support frame has an upper portion which is attached to an upper portion of the nacelle. The auxiliary support frame has a lower portion which is detachable from the upper portion of the auxiliary frame. The lower portion of the auxiliary frame is connected to a lower portion of the nacelle, the lower portion of the nacelle being detachable from the upper portion of the nacelle. By removing the lower portion of the nacelle and the lower portion of the auxiliary support frame, the thrust producing section of the turbine engine can be removed intact.

CROSS-REFERENCE 
Reference is made to a related copending application, herein incorporated 
by reference, entitled "Structure for Eliminating Lift Load Bending in 
Engine Core of Turbofan", filed Sep. 26, 1991, having U.S. patent 
application Ser. No. 765,804, (13DV-10150). 
BACKGROUND OF THE INVENTION 
The present invention relates to gas turbine engines for aircraft and, more 
particularly, to the engine housing of a gas turbine engine in which the 
fan outer nacelle and auxiliary support frame are each segmented into two 
180 degree segments to allow for easy removal and installation of the 
thrust producing section of a large turbine engine. 
Gas turbine engines for commercial aircraft generally have become 
increasingly larger with each generation of new aircraft. Some of the size 
increase is attributable to the need for increased power output of a 
turbofan engine in order to power larger aircraft. Some size increase is 
also attributable to the desire for more fuel-efficient engines achievable 
by increasing the engine bypass ratio. A higher bypass ratio requires a 
larger fan and a correspondingly larger diameter nacelle structure. 
Present generation aircraft gas turbine engines with high bypass ratios 
may have nacelle diameters in the range of twelve feet. Not only does this 
large engine create a clearance concern for under-wing mounting, but it 
also presents a concern for installation and removal on a timely basis for 
service. 
U.S. patent application Ser. No. 765,804 entitled "Structure for 
Eliminating Lift Load Bending in Engine Core of Turbofan" filed Sep. 26, 
1991, discloses a novel nacelle and frame mounting structure which is 
designed to prevent the engine core in a turbofan engine from experiencing 
large bending loads generated by inlet lift loads. This novel nacelle 
structure comprises a load carrying nacelle for efficient load transfer. 
The nacelle structure is capable of transferring loads directly from the 
inlet to the attaching pylon and isolating the engine gas generator from 
the load path. For such a nacelle structure, engine service requires 
detachment of the nacelle from the engine frame mount and removal of the 
fan blades. The outlet guide vane (OGV) assembly must also be removed 
after which the gas generator can be floated away to the aft of the 
nacelle structure. 
Since time is of the essence when engine replacement is necessitated 
(usually overnight placement being desired), it is important that measures 
be taken to insure that quick engine replacement and maintenance are 
realized. Therefore, a need exists for a nacelle structure which would 
make the easy maintenance and replacement of large turbofan engines 
possible, including those engines which have a nacelle structure designed 
to prevent core bending of the engine. 
SUMMARY OF THE INVENTION 
Accordingly, one object of the present invention is to provide a structural 
support configuration for an aircraft turbine engine which minimizes 
installation and removal time. 
Another object of the present invention is to provide a structural support 
configuration which allows the major thrust producing portion of the 
engine to remain intact when the removal of the turbine engine is desired. 
These and other objects and advantages of the present invention are 
provided by a detachable support structure for a gas turbine aircraft 
engine having a fan casing which encloses a plurality of fan blades. A fan 
OGV assembly is connected to the aft end portion of the fan casing. A 
nacelle is connected to the aircraft pylon and auxiliary frame while 
surrounding the fan casing and fan OGV assembly. The nacelle has an upper 
portion and a lower portion. The lower portion of the nacelle is 
detachable from the upper portion of the nacelle. The forward region of 
the nacelle, which includes both lower and upper portions, is 
circumferentially connected to an air inlet. 
An auxiliary support frame is connected to the nacelle at a location to the 
aft of the fan OGV assembly. The auxiliary support frame has an upper 
portion and a lower portion. The lower portion of the auxiliary support 
frame is detachable from the upper portion of the auxiliary support frame. 
The upper portion of the nacelle is connected to the aircraft pylon and 
the upper portion of the auxiliary support frame and the lower portion of 
the nacelle is connected to the lower portion of the auxiliary support 
frame. When connected, the upper and lower portions of the nacelle form a 
circle. Likewise, when connected, the upper and lower portions of the 
auxiliary frame form a second circle which is located radially inward from 
the circle formed by the connection of the upper and lower portions of the 
nacelle. 
By detaching the lower portion of the nacelle from the upper portion of the 
nacelle, and by detaching the lower portion of the auxiliary support frame 
from the upper portion of the auxiliary support frame, the fan casing, the 
plurality of fan blades enclosed by the fan casing, the fan OGV assembly, 
and remainder of turbomachinery can be removed intact. This removal 
procedure includes an initial step of insuring that the air inlet is 
removed or connected to the nacelle in such a manner that the removal or 
attachment of the thrust producing section of the turbine engine is not 
prevented. 
In short, the present invention allows for the easy and intact removal of 
the entire thrust producing section of the turbine engine.

DETAILED DESCRIPTION OF THE INVENTION 
With reference to FIG. 1, fan outer nacelle 10 is located radially outward 
from fan casing 12 and connected thereto by slip joint 14 (slip joint 14 
being designed to form a radial load path 14A). The forward section of fan 
outer nacelle 10 is comprised of inlet 15 which provides for air direction 
inward to the engine and outward to the nacelle. The rear of fan outer 
nacelle 10 connects to an auxiliary support frame 16 which connects to the 
floor 20A of pylon 20 by means of an auxiliary frame mount (to be 
discussed later). The location where the pylon floor 20A attaches to the 
auxiliary frame mount is indicated by auxiliary frame mount location 18. 
Pylon 20 extends from aircraft wing 34 and provides a means by which the 
turbine engine assembly can be attached to the aircraft. Pylon 20 extends 
from forward mount location 26 to rear mount location 32. The top of the 
pylon extends forward along the upper nacelle section. 
Continuing in FIG. fan OGV's (outlet guide vanes), which comprise a fan OGV 
assembly 22, are located directly forward of forward mount location 26 and 
connect to fan casing 12 at a location proximate to slip joint 14. Located 
radially inward from fan casing 12 are fan blades 28 which are connected 
to fan disk 30. The engine core area 36 is located between forward mount 
location 26 and rear mount location 32, the rear portion 38 of the engine 
being located to the aft of rear mount location 32. 
With reference to FIG. 2, line 40 is an axial reference which extends 
horizontally through the center of the engine. Nacelle 10 is divided into 
an upper nacelle 10A and a lower nacelle 10B. In a like manner, inlet 15 
is divided into an upper inlet 15A and a lower inlet 15B. Section 42 
indicates the area where auxiliary support frame 16 is located, and line 
44 indicates a line of demarcation by which the upper and lower halves of 
nacelle 10 and inlet 15 are divided. 
In FIG. 3, fan blades 28 are surrounded by fan casing 12 which is located 
radially inward of the circle formed by upper half nacelle 10A and lower 
half nacelle 10B. Split locations 44A and 44B separate upper half nacelle 
10A and lower half nacelle 10B into two 180 degree segments. 
With reference to FIGS. 4A and 4B, fan casing 12 (FIG. 4A) lies in the 
vertical plane indicated by line A--A of FIG. 2 and surrounds fan blades 
28. Located behind fan casing 12 is engine forward mount 52 (FIG. 4B) 
which attaches to pylon 20 at forward mount location 26 of FIG. 1. (A rear 
mount, not shown, similar to forward mount 52 connects to pylon 20 at rear 
mount location 32 of FIG. 1). Upper half auxiliary support frame 16A (FIG. 
4B) and lower half auxiliary support frame 16B (FIG. 4A) are located 
behind engine forward mount 52 and are at the auxiliary mount location 18 
of FIG. 1. Upper half auxiliary support frame 16A connects to upper 
nacelle 10A by means of struts 50A (FIGS. 4B and 5) and lower half 
auxiliary support frame 16B (FIGS. 4A and 5) connected to lower nacelle 
10B by means of struts 50B. 
The upper half auxiliary support frame 16A is attached to an auxiliary 
mount (not shown) similar to engine forward mount 52. The auxiliary mount 
attaches to pylon 20. In such a manner, upper half auxiliary support frame 
16A and upper half nacelle 10A are firmly secured to the pylon 20. 
With reference to FIG. 5, upper half auxiliary support frame 16A is 
connected to detachable lower half auxiliary support frame 16B. Upper half 
auxiliary support frame 16A is attached to auxiliary frame mount 53 which 
is secured to pylon 20. Upper half auxiliary support frame 16A is 
connected to upper nacelle 10A by means of struts 50A, and lower half 
auxiliary support frame 16B is connected to lower nacelle 10B by means of 
struts 50B. 
Since the easy removal of turbine engine components is desired, fan inlet 
15, nacelle 10, and auxiliary support frame 16 are formed by two 180 
degree sections about the 3 and 9 o'clock positions. The respective upper 
and lower sections are connected by securing means (not shown) which 
comprise clamped connections. By disconnecting lower inlet 15B from upper 
inlet 15A, by disconnecting lower nacelle 10B from upper nacelle 10A, and 
by disconnecting lower auxiliary support frame 16B from upper support 
frame 16A, the engine can be removed. Thus, the entire thrust producing 
fan section (blades, casing, and OGV's) of the engine can be removed with 
the remainder of the engine intact without the need for a laborious 
piece-by-piece disassembly. 
Further, the removed parts (blades, casing, and OGV's) could be replaced by 
an intact thrust producing section, i.e., blades 28, fan casing 12, and 
fan OGV's 22, Which is tested and ready for service. 
When it is desired to remove the thrust producing section of the engine, 
the upper nacelle 10A and lower nacelle 10B must be unbolted. However, 
since the upper nacelle 10A is connected to upper half auxiliary support 
frame 16A which is connected to lower half auxiliary support frame 16B 
(FIG. 5), it is also necessary to disconnect the lower half auxiliary 
support frame 16B from the upper half auxiliary support frame 16A. The 
lower half support frame 16B and lower half nacelle 10B can then be 
removed in the direction indicated by arrow 48 (FIG. 4A). Once the lower 
half nacelle 10B and lower half auxiliary support frame 16B are removed, 
the fan casing 12 and fan blades 28 may be removed as well (FIG. 4B). 
Since fan OGV's 22 are connected to the nacelle 10 by means of slip joints 
14 (FIG. 1) which connect to the end portion of fan casing 12, the fan 
OGV's are simply removed with the engine structure. 
It may be desired for structural and flowpath reasons to have a 360 degree 
inlet 15 such as that of the alternative embodiment of the invention 
depicted in FIG. 6 in which inlet 15 is detachably connected to nacelle 
15. If such is the case, the removal of the engine thrust producing 
section, i.e., blades, fan casing, and OGV's, would proceed as previously 
described except that the 360 degree inlet could remain supported via 
nacelle 10A or could be removed first. 
The present invention provides a configuration which minimizes engine 
installation and removal time while maintaining intact the major thrust 
producing section of the engine. 
The foregoing detailed description of the preferred embodiments of the 
invention is intended to be illustrative and non-limiting. Many changes 
and modifications are possible in light of the above teachings. Thus, it 
is understood that the invention may be practiced otherwise than as 
specifically described herein and still be within the scope of the 
appended claims.