Gas turbine engine variable bleed pivotal flow splitter

The present invention provides a gas turbine engine with a circumferentially disposed plurality of pivotal flow splitters between compressor sections to bleed off flow and remove particles, particularly ice in the case of an aircraft fan-jet gas turbine engine, by pivoting the leading edge of the splitter into the compressor flow thereby using the total pressure Q of the flow to drive bleed flows overboard and remove particles. In the preferred embodiment the pivotal flow splitter is integral with a booster variable bleed valve (VBV) door, and includes a booster bleed duct means having a bellmouth shaped inducer type inlet to enhance the capture of compressor flow and particle removal

BACKGROUND OF THE INVENTION 
1. Field of the Invention 
This invention relates to gas turbine engines variable bypasses and 
particle removal and more particularly to variable bypass flow splitters 
integral with surge bleed doors that pivot into the engine flowpath in a 
transition section between the booster and core engine compressor 
sections. 
2. Description of Related Art 
It is well known in the gas turbine engine field to provide variable bleed 
valves (VBVs) having doors that open to provide a bleed path to bleed off 
compressed air between the booster and core engine compressor of gas 
turbine engines. Aircraft fan jet gas turbine engines and marine and 
industrial derivatives of such engines have employed various forms of 
curved flowpaths and VBV bleed doors that are retracted into the flowpath 
casing so as to form an entrance to a bleed duct that bleeds booster or 
low pressure compressor discharge airflow to draw particles out of the 
flowpath in a manner such as that disclosed in U.S. Pat. No. 4,463,552 
entitled "Combined Surge Bleed and Dust Removal System for a Fan-Jet 
Engine" by Monhardt et. al. The problem with such systems is that amount 
and force of the bleed flow is dependent on the static pressure of the 
compressor airflow and is often not strong enough to remove larger pieces 
and amounts of particles such as ice. Because the bleed flow abruptly 
curves away from the direction of the compressor flow it is very difficult 
to hold larger particles in the bleed flow because of their momentum. This 
problem is common to aircraft, marine, and ground based gas turbine 
engines. Fan jet engines such as the General Electric CF6-80 series of 
engines have in series relationship a fan, a booster, and a core engine 
compressor whereby a portion of the air passing through the fan is ducted 
to the booster and then the core engine compressor. In order to match the 
inlet airflow of the core engine compressor to its flight operational 
requirements and to prevent booster stall a booster variable bleed valve 
(VBV) is provided in the form of a booster bleed duct having an inlet 
between the booster and the core engine compressor and an outlet to the 
fan duct. Opening and closing of the booster bleed duct is conventionally 
provided by a circumferentially disposed plurality of pivotal doors that 
retract into the engine structure or casing and are operated by a single 
unison ring powered by one or more fuel powered actuators. Bellcrank 
linkages operably connect the retracting pivotal bleed doors to the unison 
ring. An example of such a stall prevention system using a retracting 
pivotal door, as compared to a sliding door or valve in the Monhardt 
patent, is disclosed in U.S. Pat. No. 3,638,428 entitled "Bypass Valve 
Mechanism" by Shipley et. al. and assigned to the same assignee as the 
present invention and incorporated herein by reference The operation of 
the VBV is scheduled by the engine controller, either a mechanical or 
digital electronic type may be used. 
The problem associated with conventional bleed valve ducts and valve doors 
is that larger particles and amounts of particles such as ice are often 
not drawn into the bleed duct. The present invention provides the ability 
to remove larger particles in greater amounts from the compressor airflow 
in a more efficient manner than has been previously possible 
Modern aircraft employ fewer of the higher thrust, fuel efficient, 
very-high-bypass engines such as the twin engine Boeing 767 aircraft. 
Aircraft with fewer engines require more total take-off power and more 
power per engine in order to satisfy the requirement of being able to fly 
with one engine out. Therefore, the engines are set to lower power 
settings resulting in less engine airflow during descent when all engines 
are operational. This results in high water content for engine airflow 
since the amount of hail or water that gets into the engine is the same so 
long as the aircraft speed remains the same. On the other hand, higher 
bypass ratio engines have smaller core flow and larger bullet-nose frontal 
area. This means more hail or water gets through the compressor into the 
combustor resulting in higher water content for the air. These two 
fundamental phenomena combine to cause substantial increase of 
water-to-air ratio in the combustor resulting in such aircraft engines 
being more susceptible to engine flame out problem in rain or hail storms. 
Furthermore modern high bypass ratio engines incorporating higher pressure 
core compressors and lower pressure boosters produce less pressure 
difference between the booster exit and the fan bypass duct. Therefore, it 
increases the difficulties of bleeding sufficient amounts of air from 
downstream of the booster to the fan bypass duct for protecting boosters 
from stall. The current invention fully utilizes the dynamic pressure head 
to increase bleed capability. 
Industrial and marine derivatives of aircraft engines, generically referred 
to as ground based gas turbine engines or derivative engines, generally 
replace propulsive elements such as the fan and exhaust nozzle of an 
aircraft gas turbine engine with a load means such as an electrical 
generator, ship or marine propeller, or pump (e.g. natural gas pump). 
Electrical generator gas turbine engines are required to operate at a 
constant RPM such as 3000 RPM or 3600 RPM in order to generate electricity 
at a constant 50 Hz or 60 Hz respectively so as to be synchronized with 
the electrical grid network for which they are used to supply electricity. 
Two general types of derivative engines are free turbine and direct drive 
engines. Free turbine types employ a free power turbine that drives the 
load and is not directly mechanically coupled to a compressor (usually the 
booster or low pressure compressor of a multiple rotor engine) of the gas 
generator used to power the free power turbine. The direct drive turbine 
is directly mechanically coupled to both a compressor of the gas turbine 
engine and to the load on a common rotor that is most often the low 
pressure rotor of a dual rotor gas turbine engine. 
A difficult problem is bringing a multiple rotor direct drive engine system 
to a no-load synchronous speed condition prior to locking in on an 
electrical grid network which would then hold the low pressure (LP) rotor 
attached system in a synchronous speed mode. The problem is how to reach 
zero shaft horsepower (SHP) output with the LP system at synchronous speed 
without stalling the booster compressor so as to connect and disconnect 
the gas turbine driven generator from the electrical grid network while 
not overspeeding the LP system. 
The booster stall margin must be controlled by controlling either the inlet 
flow of the booster or by controlling the booster discharge flow level 
entering the compressor. Typically booster discharge bleed doors are 
opened to dump some of the booster flow overboard so as to control the 
booster operating line to a point below its stall line. Aircraft engine 
VBV doors normally are inadequate to allow such operation even with 
state-of-the-art variable inlet guide vane (VIGV) closure on the booster. 
Current VBV door sizes can be up to 2 times too small to reach this 
condition. 
In adapting an aircraft engine as a derivative for this type of 
application, since LP speed is held constant to very low powers, a means 
is needed to minimize required engine changes to accomplish a no-load 
synchronous speed. For example, it is estimated that for an industrial 
derivative of the CF6-80C2 engine its VBV bleed doors wide open will only 
flow 50% of the bleed flow needed to set the booster operating line no 
higher than its maximum allowable pressure ratio. For example, either the 
VBV flow area must be doubled or the booster inlet flow must bereduced 
from 145 lb/sec down to 105 lb/sec. 
In accomplishing this, the inlet flow to the core engine compressor must 
not suffer increased pressure or temperature distortion which could stall 
the compressor and possibly harm the engine. It is well known that 
variable inlet guide vanes on boosters and compressors may be used to 
control the amount of inlet flow and boost pressure at a constant rotor 
speed by their changing of incidence flow angles to the rotors but it is 
very difficult if not impossible to avoid stalling the booster or low 
compressor using this method. 
SUMMARY OF THE INVENTION 
The present invention provides a gas turbine engine with a 
circumferentially disposed plurality of pivotal flow splitters and a means 
to pivot the leading edges of the splitters into the compressor flowpath 
thereby fully using the dynamic pressure head (Q) of the flow to maximize 
the bleed air flowrate and improve particle removal. 
One embodiment provides an aircraft gas turbine engine having a VBV between 
a booster and core engine compressor sections including a flow splitter 
integral with a VBV door such that the splitter's leading edge can be 
pivoted into the compressor flowpath between two compressor sections of 
two respective rotors. This apparatus uses Q to enhance the particle 
removal process to remove dirt, ice, and other particles from the 
compressor flowpath of a gas turbine engine and to enhance the VBV 
function by increasing the flowrate. 
The preferred embodiment provides a boundary layer suction means in the 
form of sealable boundary layer flow leakage paths along the hinge line 
and sides of the pivotal splitter or door that are exposed when the door 
is opened and covered when the door is closed. This reduces the 
possibility and degree of boundary layer separation and stalling of the 
core compressor downstream of the splitter. 
The preferred embodiment also provides a controlled diffusion surface on 
the engine flow side of the door that is effective to prevent flow 
separation when the door is fully deployed thereby preventing distorted 
flow from entering and possibly even stalling the core compressor 
downstream of the splitter. 
The preferred embodiment includes sidewalls along the sides of the door 
forming a chute and a roof attached to the top of the chute walls over the 
door to form a scoop. In the preferred embodiment the scoop is in the form 
of a duct that is integral with the booster variable bleed valve (VBV) 
door and the booster bleed duct means includes an inducer type inlet in 
having a bellmouth inlet duct operably positioned to receive bleed flow 
from the scoop and capture the bleed flow's dynamic pressure. 
The present invention has the advantage of using the dynamic pressure of 
the engine flowpath in order to more effectively remove particles from the 
engine flowpath than particle removal apparatus found in the prior art. 
Another advantage of the present invention is that it allows the use of an 
inducer in the bleed duct means to further enhance the particle removal 
process of the gas turbine engine. 
Another embodiment of the present invention is for a ground based gas 
turbine engine, preferably a derivative of an aircraft gas turbine engine. 
The preferred embodiment employs a dual rotor, high pressure and low 
pressure, direct drive derivative gas turbine engine powering an 
electrical generator wherein the generator is directly coupled to the low 
pressure, or booster, compressor section. The engine's VBV, located 
between booster and core engine compressor sections is provided with a 
pivoting flow splitter integral with a VBV door such that the splitter's 
leading edge can be pivoted into the compressor flowpath between two 
compressor sections of two respective rotors. 
This apparatus uses Q to enhance the VBV's flow bleed operation so as to 
provide a way to bring the electrical generator to a no-load synchronous 
speed condition prior to locking in on an electrical grid network without 
stalling the booster compressor. The invention also provides a means to 
bring the electrical generator to a no-load synchronous speed condition 
prior to disconnecting the generator from the electrical grid network 
without stalling the booster compressor including a quick full power fuel 
stopcock, wherein conventionally, IGV's are slammed close and VBV's are 
slammed open.

DETAILED DESCRIPTION OF THE INVENTION 
A first embodiment of the present invention is illustrated in FIG. 1 for a 
fan-jet gas turbine engine (not fully shown) having a fan 8 that 
pressurizes and feeds air to a fan bleed duct 10 disposed between an inner 
fan case 11 and an outer fan case 12 and to an engine coreflowpath 13. A 
booster 20 located at a forward portion of engine core flowpath 13 rotates 
together with fan 8 on a low pressure rotor 22 driven by a conventional 
low pressure turbine (not shown) and further compresses air that is then 
ducted to a conventional engine core compressor 24. 
A circumferentially disposed plurality of booster bleed paths 16 indicated 
by the arrows and dotted line labeled 16 are radially disposed between 
engine core flowpath 13 and fan bleed duct 10 employing fan struts 14 as 
sidewalls and a radially outwardly disposed bleed duct 18. A booster bleed 
inlet 26, an opening, is disposed in the wall of a transition duct 27 
between booster 20 and engine core compressor 24 in engine core flowpath 
13 and is controlled by the engine's VBV system as illustrated by a VBV 
door 30 which serves as a pivotal flow splitter having a leading edge 31 
that controls and schedules booster bleed air. 
VBV door 30 and leading edge 31 are arranged to pivot into the compressor 
flowpath between booster 20 and engine core compressor 24 to split off and 
bypass a portion of the flow in transition duct 27 and remove the 
particles entrained in that flow portion. Note that flowpath 13 through 
booster 20 is curved radially inward thus subjecting the particles 
entrained in the air flow of engine core flowpath 13 to a sudden flow 
direction change which cannot sufficiently overcome the larger momentum of 
the particles. Therefore the particles are intercepted by VBV door 30 and 
directed into bleed path 16. 
Engine core flowpath 13 is curved radially inward through booster 20 and 
radially outward through engine core compressor section 24 in a quasi 
sinusoidal shape defining a transition section 27 between booster 20 and 
core compressor section 24. Engine core flowpath 13 through transition 
section 27 has aerodynamically curved sidewalls that smoothly expand the 
airflow therethrough to prevent flow separation prior to entry to engine 
core compressor 24. A controlled diffusion surface 33 is provided on the 
transition duct side of VBV door 30 such that surface 33 is 
aerodynamically curved so as to prevent separation of the flow along the 
surface. 
As illustrated in FIG. 1, pivotal VBV door 30 is disposed in booster bleed 
inlet 26 and is operable to pivot through a variable angle A, defining the 
fully open and closed positions of VBV door 30, in order to vary the 
amount of flow passing through booster bleed paths 16 that are 
circumferentially disposed around the engine. Angle A and hence the flow 
rate of booster bleed air is controlled by a VBV control system (not 
shown) as is conventionally known and used in the art. Opening and closing 
of VBV doors 30 is conventionally provided by bellcrank linkages 40, 
located in a chamber generally shown at 44, that are actuated by a single 
unison ring powered by one or more fuel powered actuators (not shown). 
Surface 33 is preferably designed to prevent separation at a maximum angle 
A, which has been found to preferably be about 20 degrees. 
Referring to FIGS. 1 and 2, VBV door 30 and its leading edge 31 pivots into 
engine core flowpath 13 and includes a scoop 32 having a relatively wide 
scoop inlet 34 converging to a relatively narrow scoop outlet 36 and a 
contoured converging scoop passage 38 therebetween. 
Scoop inlet 34 including leading edge 31 faces into the flow of the engine 
so as to scoop up both engine core airflow and particles entrained in the 
portion of the flow crossing scoop inlet 34. Scoop 32 and its contoured 
contracting scoop passage 38 is designed to enhance the flow through 
booster bleed paths 16 and into bleed duct 18. To further enhance this 
flow, chamber 44 includes a bellmouth shaped inducer inlet 48 wherein the 
cross-section of the opening of inducer inlet 48 is of a size and shape 
generally corresponding to that of scoop outlet 36 but is appreciably 
larger. 
Note that an actuating rod 41 of linkage 40 is attached to the top of scoop 
32 thereby not producing any aerodynamic interference with the operation 
of scoop 32. 
A boundary layer suction means, depicted in FIG. 2 as sealable boundary 
layer flow leakage paths along VBV door hinge line 52 and side lines 54 of 
pivotal door 30 that are exposed when the door is opened and covered when 
the door is closed. The leakage flow, indicated by lines of leakage arrows 
along VBV door hinge line 52 and side lines 54 of pivotal door 30, reduces 
the possibility and degree of boundary layer separation and stalling of 
the core compressor downstream of the splitter. A side line shiplap seal 
56 in FIG. 2A and a hinge line shiplap seal 58 in FIG. 2B provide a 
deployable sealing means for their respective boundary layer flow leakage 
paths. A similar leading edge shiplap seal may be used in conjunction with 
leading edge 31 to help fully seal the opening, booster bleed inlet 26, 
when door 30 is in the closed position. 
Alternate embodiments to pivotal splitter 30 integral with scoop 32 having 
a contoured scoop passage 38 as illustrated in FIG. 2 are illustrated in 
FIG. 3. A chute 42 is formed by attaching sidewalls 43 to sides 45 of VBV 
door 30 from aft of its leading edge 31 forming an alternate scoop passage 
38'. Scoop 32' is formed by the addition of a roof 47 to sidewalls 43 of 
chute 42 and includes a scoop inlet 34. Alternative embodiments may 
provide either a door 30 or a chute 42 alone as a pivotal splitter that 
would pivot open into core engine flow 13 so as to provide a pivotal 
splitter for booster bleed inlet 26 that faces into the engine flow and 
could capture the particles entrained in core engine flow 13 and direct 
them into bleed path 16. 
A second embodiment of the present invention is illustrated in FIG. 4 for a 
derivative gas turbine engine 2 of an aircraft gas turbine engine 
illustrated in FIG. 1. An electrical generator 4 is powered by derivative 
gas turbine engine 2 having a booster 20 located at a forward portion of 
engine core flowpath 13. Booster 20 rotates on a low pressure rotor 22 
driven by a conventional low pressure turbine 21 that also drives 
electrical generator 4 directly coupled to low pressure rotor 22. A core 
engine compressor 24 is located downstream of booster 20 and is driven by 
a core engine turbine 29 through a high pressure rotor 35. 
A circumferentially disposed plurality of booster bleed valves 25 having 
booster bleed inlets 26 are disposed in a wall of a transition duct 27 
between booster 20 and engine core compressor 24 in engine core flowpath 
13. VBV 25 is controlled by the engine's VBV system as illustrated by a 
VBV door 30 which serves as a pivotal flow splitter having a leading edge 
31. VBV 25 controls and schedules booster bleed air for derivative gas 
turbine engine 2. 
VBV door 30 and leading edge 31 are arranged to be pivoted into the 
compressor flowpath between booster 20 and engine core compressor 24 in a 
manner similar to that of the aircraft fan-jet gas turbine engine 
illustrated in FIG. 1. However, note that VBV door 30 for a derivative gas 
turbine engine 2 as illustrated herein is preferably longer than one for 
an aircraft gas turbine engine. This provides for bleeding or bypassing a 
larger amount of flow while retaining a relatively small door opening 
variable angle A to prevent separation along a controlled diffusion 
surface 33' of door 30. 
Inducers similar to those illustrated by bellmouth shaped inducer inlet 48 
in FIGS. 1 and 2 may also be used downstream of VBV 30 in the flow along 
booster bleed path 16 to enhance the bleed process as described above and 
illustrated in FIG. 1. 
The invention has been described for use in a fan-jet aircraft gas turbine 
engine and a derivative gas turbine engine powering an electrical 
generator. However, it is contemplated for use in other gas turbine 
engines including other derivative engines such as marine and industrial 
gas turbine engines. Also contemplated are engines having more than two 
rotors. 
Scheduling the opening of VBV door 30 to its maximum position is preferably 
a function dependent on one or more engine operating conditions. It has 
been found preferable to calibrate the door's opening schedule as a 
function of booster speeds and core engine speeds for a dual rotor direct 
drive derivative engine as depicted in FIG. 4. More particularly, the door 
opening schedule could be a function of the following parameters, booster 
corrected speed (N.sub.2 /.sqroot..THETA..sub.2) or core speed corrected 
to the core compressor inlet conditions (N.sub.25 
/.sqroot..THETA..sub.25), whereby: 
N.sub.2 is the physical or directly measured rotational speed of low 
pressure rotor 22, 
(.THETA..sub.2) is the ratio of the booster 20 inlet temperature T.sub.2 
.degree.R/519.degree. R. and 519.degree. R. corresponds to an industry 
standard 59.degree. F. day, 
N.sub.25 is the physical or directly measured rotational speed of high 
pressure rotor 35, and 
(.THETA..sub.25) is the ratio of core engine compressor 24 inlet 
temperature T.sub.25 .degree.R/519.degree. R. 
While the preferred embodiment of the present invention has been described 
fully in order to explain its principles, it is understood that various 
modifications or alterations may be made to the preferred embodiment 
without departing from the scope of the invention as set forth in the 
appended claims.