U.S. patent application number 16/105220 was filed with the patent office on 2018-12-13 for geared gas turbine engine with reduced oil tank size.
The applicant listed for this patent is United Technologies Corporation. Invention is credited to Michael E. McCune, William G. Sheridan.
Application Number | 20180355803 16/105220 |
Document ID | / |
Family ID | 54196539 |
Filed Date | 2018-12-13 |
United States Patent
Application |
20180355803 |
Kind Code |
A1 |
Sheridan; William G. ; et
al. |
December 13, 2018 |
GEARED GAS TURBINE ENGINE WITH REDUCED OIL TANK SIZE
Abstract
A gas turbine engine comprises a fan drive turbine for driving a
gear reduction, which drives a fan rotor. A lubrication system
supplies oil to the gear reduction. An oil tank is relatively
small. The lubrication system operates to allow oil to remain in
the oil tank for a dwell time of less than or equal to five
seconds. A method of designing a gas turbine engine is also
disclosed.
Inventors: |
Sheridan; William G.;
(Southington, CT) ; McCune; Michael E.;
(Colchester, CT) |
|
Applicant: |
Name |
City |
State |
Country |
Type |
United Technologies Corporation |
Farmington |
CT |
US |
|
|
Family ID: |
54196539 |
Appl. No.: |
16/105220 |
Filed: |
August 20, 2018 |
Related U.S. Patent Documents
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Application
Number |
Filing Date |
Patent Number |
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14595255 |
Jan 13, 2015 |
10054058 |
|
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16105220 |
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61929174 |
Jan 20, 2014 |
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Current U.S.
Class: |
1/1 |
Current CPC
Class: |
F16H 57/0486 20130101;
F01D 25/20 20130101; F16H 57/0479 20130101; F02C 7/36 20130101;
F05D 2220/323 20130101; F16H 57/0427 20130101; F02K 3/06 20130101;
F05D 2260/40311 20130101; F05D 2260/608 20130101; F16H 57/04
20130101; F02C 7/06 20130101; Y02T 50/60 20130101; F16H 57/08
20130101; F05D 2260/98 20130101; Y02T 50/671 20130101 |
International
Class: |
F02C 7/36 20060101
F02C007/36; F16H 57/04 20060101 F16H057/04; F02K 3/06 20060101
F02K003/06; F01D 25/20 20060101 F01D025/20 |
Claims
1. A gas turbine engine comprising: a fan drive turbine driving a
gear reduction, said gear reduction driving a fan rotor, and said
gear reduction including a sun gear driving intermediate gears; a
lubrication system supplying oil to said gear reduction and
comprising an oil tank, wherein the oil remains in the oil tank for
a dwell time of less than or equal to five seconds during
operation; an oil capture gutter surrounding said gear reduction;
and a gear ratio of said gear reduction greater than or equal to
2.3:1.
2. The gas turbine engine as set forth in claim 1, wherein said
dwell time is less than or equal to 3.0 seconds during
operation.
3. The gas turbine engine as set forth in claim 2, further
comprising oil baffles located circumferentially between said
intermediate gears.
4. The gas turbine engine as set forth in claim 3, wherein said fan
rotor delivers air into a bypass duct as propulsion air and into a
core engine where it passes into a compressor section, with a
bypass ratio defined as the ratio of air delivered into said bypass
duct compared to the volume of air delivered into said core engine,
said bypass ratio being greater than or equal to 10.0.
5. The gas turbine engine as set forth in claim 4, wherein said oil
tank holds greater than or equal to 25 and less than or equal to 35
quarts of oil.
6. The gas turbine engine as set forth in claim 4, wherein said oil
tank holds greater than or equal to 35 and less than or equal to 50
quarts of oil.
7. The gas turbine engine as set forth in claim 6, wherein said oil
tank is associated with an engine having greater than or equal to
35,000 and less than or equal to 100,000 lbs in rated thrust at
take-off.
8. The gas turbine engine as set forth in claim 7, wherein the fan
drive turbine includes an inlet, an outlet, and a fan drive turbine
pressure ratio greater than 5, wherein the fan drive turbine
pressure ratio is a ratio of a pressure measured prior to the inlet
as related to a pressure at the outlet prior to any exhaust nozzle,
and the fan rotor includes a plurality of fan blades, a fan
pressure ratio across the fan blades of less than 1.45, measured
across the fan blades alone.
9. The gas turbine engine as set forth in claim 1, wherein said fan
rotor delivers air into a bypass duct as propulsion air and into a
core engine where it passes into a compressor section, with a
bypass ratio defined as the ratio of air delivered into said bypass
duct compared to the volume of air delivered into said core engine
with said bypass ratio being greater than or equal to 10.0.
10. The gas turbine engine as set forth in claim 9, wherein said
oil tank holds greater than or equal to 25 and less than or equal
to 35 quarts of oil.
11. The gas turbine engine as set forth in claim 9, wherein said
oil tank holds greater than or equal to 35 and less than or equal
to 50 quarts of oil.
12. A gas turbine engine comprising: a fan drive turbine driving a
gear reduction, said gear reduction driving a fan rotor, and said
gear reduction including a sun gear driving intermediate gears; a
lubrication system supplying oil to said gear reduction and
comprising an oil tank, wherein the oil remains in the oil tank for
a dwell time; a gear ratio of said gear reduction greater than or
equal to 2.3:1; and wherein said dwell time is less than or equal
to 3.0 seconds during operation.
13. The gas turbine engine as set forth in claim 12, further
comprising oil baffles located circumferentially between said
intermediate gears.
14. The gas turbine engine as set forth in claim 13, further
comprising an oil capture gutter surrounding said gear
reduction.
15. The gas turbine engine as set forth in claim 14, wherein said
oil tank holds greater than or equal to 25 and less than or equal
to 35 quarts of oil.
16. The gas turbine engine as set forth in claim 15, wherein said
engine is rated greater than or equal to 15,000 and less than or
equal to 35,000 lbs in rated thrust at take-off.
17. The gas turbine engine as set forth in claim 12, wherein said
oil tank holds greater than or equal to 25 and less than or equal
to 35 quarts of oil, and said engine is rated greater than or equal
to 15,000 and less than or equal to 35,000 lbs in rated thrust at
take-off.
18. The gas turbine engine as set forth in claim 17, wherein the
fan drive turbine includes an inlet, an outlet, and a fan drive
turbine pressure ratio greater than 5, wherein the fan drive
turbine pressure ratio is a ratio of a pressure measured prior to
the inlet as related to a pressure at the outlet prior to any
exhaust nozzle, and the fan rotor includes a plurality of fan
blades, a fan pressure ratio across the fan blades of less than
1.45, measured across the fan blades alone.
19. The gas turbine engine as set forth in claim 12, wherein the
fan drive turbine includes an inlet, an outlet, and a fan drive
turbine pressure ratio greater than 5, wherein the fan drive
turbine pressure ratio is a ratio of a pressure measured prior to
the inlet as related to a pressure at the outlet prior to any
exhaust nozzle, and the fan rotor includes a plurality of fan
blades, a fan pressure ratio across the fan blades of less than
1.45, measured across the fan blades alone.
20. The gas turbine engine as set forth in claim 19, wherein the
fan blades have a fan tip speed of less than 1150 ft/second, and
the gear reduction is a planetary gear system.
21. The gas turbine engine as set forth in claim 12, wherein an oil
capture gutter surrounds said gear reduction, and said gear
reduction is a planetary gear system.
22. A method of designing a gas turbine engine comprising:
providing a fan drive turbine driving a gear reduction, said gear
reduction driving a fan rotor; providing a lubrication system
supplying oil to said gear reduction, wherein the oil remains in
the oil tank for a dwell time of less than or equal to five seconds
during operation, and said gear reduction comprises a gear ratio
greater than or equal to 2.3:1 and includes a sun gear that drives
intermediate gears; and said fan rotor delivering air into a bypass
duct as propulsion air and into a core engine where it passes into
a compressor section, with a bypass ratio defined as the ratio of
air delivered into said bypass duct compared to the volume of air
delivered into said core engine, said bypass ratio greater than or
equal to 10.0.
23. The method as set forth in claim 22, wherein said dwell time is
less than or equal to 3.0 seconds during operation.
24. The method as set forth in claim 23, further comprising oil
baffles circumferentially between said intermediate gears.
25. The method as set forth in claim 24, further comprising an oil
capture gutter around said gear reduction.
26. The method as set forth in claim 25, wherein said oil tank
holds greater than or equal to 35 and less than or equal to 50
quarts of oil, and said oil tank is associated with an engine
having greater than or equal to 35,000 and less than or equal to
100,000 lbs in rated thrust at take-off.
27. The method as set forth in claim 26, wherein the fan drive
turbine includes an inlet, an outlet, and a fan drive turbine
pressure ratio greater than 5, wherein the fan drive turbine
pressure ratio is a ratio of a pressure measured prior to the inlet
as related to a pressure at the outlet prior to any exhaust
nozzle.
28. The method as set forth in claim 27, wherein the fan rotor
includes a plurality of fan blades, a fan pressure ratio across the
fan blades of less than 1.45, measured across the fan blades
alone.
29. The method as set forth in claim 28, wherein the fan blades
have a fan tip speed of less than 1150 ft/second.
30. The method as set forth in claim 29, further comprising a
mid-turbine frame positioned intermediate the fan drive turbine and
a second turbine.
Description
CROSS-REFERENCE TO RELATED APPLICATION
[0001] This application is a continuation of U.S. patent
application Ser. No. 14/595,255 filed Jan. 13, 2015, which claims
priority to U.S. Provisional Patent Application No. 61/929,174,
filed Jan. 20, 2014.
BACKGROUND OF THE INVENTION
[0002] This application relates to a gas turbine engine having a
gear reduction driving a fan wherein an oil tank is of reduced
size.
[0003] Gas turbine engines are known and, typically, include a fan
delivering air into a bypass duct as propulsion air. The fan also
delivers air into a core engine where it passes to a compressor.
The air is compressed in the compressor and delivered downstream
into a combustion section where it is mixed with fuel and ignited.
Products of this combustion pass downstream over turbine rotors
driving them to rotate.
[0004] Historically, the fan rotor and a fan drive turbine rotor
have been driven at the same speed. This placed a restriction on
the desirable speed of both the fan and the fan drive turbine.
[0005] More recently, it has been proposed to provide a gear
reduction between the fan drive turbine and the fan rotor.
[0006] The gear reduction is a source of increased heat loss. As an
example, a geared turbofan engine creates about twice as much heat
loss as a non-geared turbofan engine. In addition, the weight of
the engine increases due to the weight of the gear reduction.
[0007] It has typically been the case that a designer of a gas
turbine engine sizes an oil tank such that the oil can sit in the
oil tank long enough to de-aerate. On a normal turbofan engine,
this had been approximately at least ten seconds.
SUMMARY OF THE INVENTION
[0008] In a featured embodiment, a gas turbine engine comprises a
fan drive turbine for driving a gear reduction, which drives a fan
rotor. A lubrication system supplies oil to the gear reduction. An
oil tank is relatively small. The lubrication system operates to
allow oil to remain in the oil tank for a dwell time of less than
or equal to five seconds.
[0009] In another embodiment according to the previous embodiment,
the dwell time is less than or equal to 3.0 seconds.
[0010] In another embodiment according to any of the previous
embodiments, the gear reduction includes a sun gear for driving
intermediate gears. There are oil baffles located circumferentially
between the intermediate gears.
[0011] In another embodiment according to any of the previous
embodiments, an oil capture gutter surrounds the gear
reduction.
[0012] In another embodiment according to any of the previous
embodiments, the fan rotor delivers air into a bypass duct as
propulsion air and into a core engine where it passes into a
compressor section. A bypass ratio is defined as the ratio of air
delivered into the bypass duct compared to the volume of air
delivered into the core engine. The bypass ratio is greater than or
equal to about 6.0.
[0013] In another embodiment according to any of the previous
embodiments, the bypass ratio is greater than or equal to about
10.0.
[0014] In another embodiment according to any of the previous
embodiments, a gear ratio of the gear reduction is greater than or
equal to about 2.3:1.
[0015] In another embodiment according to any of the previous
embodiments, the oil tank may hold greater than or equal to 25 and
less than or equal to 35 quarts of oil.
[0016] In another embodiment according to any of the previous
embodiments, the engine is rated greater than or equal to 15,000
and less than or equal to 35,000 lbs in rated thrust at
take-off.
[0017] In another embodiment according to any of the previous
embodiments, the oil tank holds greater than or equal to 35 and
less than or equal to 50 quarts of oil.
[0018] In another embodiment according to any of the previous
embodiments, the oil tank is associated with an engine having
greater than or equal to 35,000 and less than or equal to 100,000
lbs in rated thrust at take-off.
[0019] In another embodiment according to any of the previous
embodiments, the gear reduction includes a sun gear for driving
intermediate gears. There are oil baffles located circumferentially
between the intermediate gears.
[0020] In another embodiment according to any of the previous
embodiments, an oil capture gutter surrounds the gear
reduction.
[0021] In another embodiment according to any of the previous
embodiments, the oil tank may hold greater than or equal to 25 and
less than or equal to 35 quarts of oil.
[0022] In another embodiment according to any of the previous
embodiments, the engine is rated greater than or equal to 15,000
and less than or equal to 35,000 lbs in rated thrust at
take-off.
[0023] In another embodiment according to any of the previous
embodiments, the oil tank holds greater than or equal to 35 and
less than or equal to 50 quarts of oil.
[0024] In another embodiment according to any of the previous
embodiments, the oil tank is associated with an engine having
greater than or equal to 35,000 and less than or equal to 100,000
lbs in rated thrust at take-off.
[0025] In another embodiment according to any of the previous
embodiments, the oil tank may hold greater than or equal to 25 and
less than or equal to 35 quarts of oil. The engine is rated greater
than or equal to 15,000 and less than or equal to 35,000 lbs in
rated thrust at take-off.
[0026] In another embodiment according to any of the previous
embodiments, the oil tank holds greater than or equal to 35 and
less than or equal to 50 quarts of oil. The oil tank is associated
with an engine having greater than or equal to 35,000 and less than
or equal to 100,000 lbs in rated thrust at take-off.
[0027] In another embodiment according to any of the previous
embodiments, an oil capture gutter surrounds the gear
reduction.
[0028] In another featured embodiment, a method of designing a gas
turbine engine comprises the steps of providing a fan drive turbine
for driving a gear reduction. The gear reduction drives a fan
rotor. A lubrication system is provided to supply oil to the gear
reduction. An oil tank is relatively small. The lubrication system
operates to allow oil to remain in the oil tank for a dwell time of
less than or equal to five seconds.
[0029] In another embodiment according to the previous embodiment,
the dwell time is less than or equal to 3.0 seconds.
[0030] In another embodiment according to any of the previous
embodiments, the gear reduction includes a sun gear for driving
intermediate gears. There are oil baffles located circumferentially
between the intermediate gears.
[0031] In another embodiment according to any of the previous
embodiments, an oil capture gutter around the gear reduction is
provided.
[0032] In another embodiment according to any of the previous
embodiments, the fan rotor delivers air into a bypass duct as
propulsion air and into a core engine where it passes into a
compressor section. A bypass ratio is defined as the ratio of air
delivered into the bypass duct compared to the volume of air
delivered into the core engine. The bypass ratio is greater than or
equal to about 10.0.
[0033] In another embodiment according to any of the previous
embodiments, a gear ratio of the gear reduction is greater than or
equal to about 2.3:1.
[0034] In another embodiment according to any of the previous
embodiments, the oil tank may hold greater than or equal to 25 and
less than or equal to 35 quarts of oil.
[0035] In another embodiment according to any of the previous
embodiments, the engine is rated greater than or equal to 15,000
and less than or equal to 35,000 lbs in rated thrust at
take-off.
[0036] In another embodiment according to any of the previous
embodiments, the oil tank holds greater than or equal to 35 and
less than or equal to 50 quarts of oil.
[0037] In another embodiment according to any of the previous
embodiments, the oil tank is associated with an engine having
greater than or equal to 35,000 and less than or equal to 100,000
lbs in rated thrust at take-off.
[0038] These and other features may be best understood from the
following drawings and specification.
BRIEF DESCRIPTION OF THE DRAWINGS
[0039] FIG. 1 is a schematic view of a gas turbine engine.
[0040] FIG. 2 shows a portion of a gear reduction.
[0041] FIG. 3 shows another portion of a gear reduction.
[0042] FIG. 4 shows a lubrication system.
DETAILED DESCRIPTION
[0043] FIG. 1 schematically illustrates a gas turbine engine 20.
The gas turbine engine 20 is disclosed herein as a two-spool
turbofan that generally incorporates a fan section 22, a compressor
section 24, a combustor section 26 and a turbine section 28.
Alternative engines might include an augmentor section (not shown)
among other systems or features. The fan section 22 drives air
along a bypass flow path B in a bypass duct defined within a
nacelle 15, while the compressor section 24 drives air along a core
flow path C for compression and communication into the combustor
section 26 then expansion through the turbine section 28. Although
depicted as a two-spool turbofan gas turbine engine in the
disclosed non-limiting embodiment, it should be understood that the
concepts described herein are not limited to use with two-spool
turbofans as the teachings may be applied to other types of turbine
engines including three-spool architectures.
[0044] The exemplary engine 20 generally includes a low speed spool
30 and a high speed spool 32 mounted for rotation about an engine
central longitudinal axis A relative to an engine static structure
36 via several bearing systems 38. It should be understood that
various bearing systems 38 at various locations may alternatively
or additionally be provided, and the location of bearing systems 38
may be varied as appropriate to the application.
[0045] The low speed spool 30 generally includes an inner shaft 40
that interconnects a fan 42, a first (or low) pressure compressor
44 and a first (or low) pressure turbine 46. The inner shaft 40 is
connected to the fan 42 through a speed change mechanism, which in
exemplary gas turbine engine 20 is illustrated as a geared
architecture 48 to drive the fan 42 at a lower speed than the low
speed spool 30. The high speed spool 32 includes an outer shaft 50
that interconnects a second (or high) pressure compressor 52 and a
second (or high) pressure turbine 54. A combustor 56 is arranged in
exemplary gas turbine 20 between the high pressure compressor 52
and the high pressure turbine 54. A mid-turbine frame 57 of the
engine static structure 36 is arranged generally between the high
pressure turbine 54 and the low pressure turbine 46. The
mid-turbine frame 57 further supports bearing systems 38 in the
turbine section 28. The inner shaft 40 and the outer shaft 50 are
concentric and rotate via bearing systems 38 about the engine
central longitudinal axis A which is collinear with their
longitudinal axes.
[0046] The core airflow is compressed by the low pressure
compressor 44 then the high pressure compressor 52, mixed and
burned with fuel in the combustor 56, then expanded over the high
pressure turbine 54 and low pressure turbine 46. The mid-turbine
frame 57 includes airfoils 59 which are in the core airflow path C.
The turbines 46, 54 rotationally drive the respective low speed
spool 30 and high speed spool 32 in response to the expansion. It
will be appreciated that each of the positions of the fan section
22, compressor section 24, combustor section 26, turbine section
28, and fan drive gear system 48 may be varied. For example, gear
system 48 may be located aft of combustor section 26 or even aft of
turbine section 28, and fan section 22 may be positioned forward or
aft of the location of gear system 48.
[0047] The engine 20 in one example is a high-bypass geared
aircraft engine. In a further example, the engine 20 bypass ratio
is greater than or equal to about six (6), with an example
embodiment being greater than about ten (10), the geared
architecture 48 is an epicyclic gear train, such as a planetary
gear system or other gear system, with a gear reduction ratio of
greater than about 2.3 and the low pressure turbine 46 has a
pressure ratio that is greater than about five. In one disclosed
embodiment, the engine 20 bypass ratio is greater than or equal to
about ten (10:1), the fan diameter is significantly larger than
that of the low pressure compressor 44, and the low pressure
turbine 46 has a pressure ratio that is greater than about five
5:1. Low pressure turbine 46 pressure ratio is pressure measured
prior to inlet of low pressure turbine 46 as related to the
pressure at the outlet of the low pressure turbine 46 prior to an
exhaust nozzle. The geared architecture 48 may be an epicycle gear
train, such as a planetary gear system or other gear system, with a
gear reduction ratio of greater than about 2.3:1. It should be
understood, however, that the above parameters are only exemplary
of one embodiment of a geared architecture engine and that the
present invention is applicable to other gas turbine engines
including direct drive turbofans.
[0048] A significant amount of thrust is provided by the bypass
flow B due to the high bypass ratio. The fan section 22 of the
engine 20 is designed for a particular flight condition--typically
cruise at about 0.8 Mach and about 35,000 feet. The flight
condition of 0.8 Mach and 35,000 ft, with the engine at its best
fuel consumption--also known as "bucket cruise Thrust Specific Fuel
Consumption (`TSFCT`)" is the industry standard parameter of lbm of
fuel being burned divided by lbf of thrust the engine produces at
that minimum point. "Low fan pressure ratio" is the pressure ratio
across the fan blade alone, without a Fan Exit Guide Vane ("FEGV")
system. The low fan pressure ratio as disclosed herein according to
one non-limiting embodiment is less than about 1.45. "Low corrected
fan tip speed"-is the actual fan tip speed in ft/sec divided by an
industry standard temperature correction of [(Tram
.degree.R)/(518.7 .degree.R)].sup.0.5. The "Low corrected fan tip
speed" as disclosed herein according to one non-limiting embodiment
is less than about 1150 ft/second.
[0049] As shown in FIG. 2, a flexible shaft 99, which is driven by
the turbine 46, drives a sun gear 101 which, in turn, engages and
drives intermediate gears 102. In some embodiments, the
intermediate gears 102 may be planet gears of a planetary epicyclic
gear system. In other embodiments, the intermediate gears 102 may
be star gears of a star epicyclic gear system. The intermediate
gears 102 engage and drive a ring gear 103 to, in turn, drive an
output shaft 106, which then drives the fan rotor 42. In other
embodiments, a planetary gear carrier (not shown) driven by
planetary gears may drive the fan shaft. Lubricant is supplied to a
journal pin 108, to the intermediate gears 102 and to other
locations within the gear reduction 48.
[0050] FIG. 3 shows baffles 100 which are placed circumferentially
between adjacent planet gears 102.
[0051] A gutter 104 surrounds the gear reduction 48 and captures
oil that has left the gear reduction. Oil from the gear reduction
48 is returned to a pump 72 (See FIG. 4) or a tank 90 as shown
schematically in FIG. 4. As shown, a lubricant system 70 includes
the gear reduction 48 which may be structured as shown in FIGS. 2
and 3. Notably, complete details of the operation of the baffle,
the gutter and the other portions of the gear reduction may be as
disclosed in U.S. Pat. No. 6,223,616, the disclosure of which with
regard to the operation of the gear reduction is incorporated by
reference.
[0052] Oil flows from an oil pump 72 to a filter 74 through a
pressure relief valve 76 to an air/oil cooler 78 and then to a
fuel/oil cooler 80. The oil may pass through an oil pressure trim
orifice 82 and back to the tank 90. Alternatively, the oil may pass
through a strainer 84 and then to the gear reduction 48. Oil
returning from the gear reduction and, in particular, from the
gutter, may pass back directly to the pump 72 or to the tank 90.
This is a simplification of the overall lubricant system and, as
appreciated, there may be other components.
[0053] Applicant has recognized that by utilizing baffles 100 and a
gutter 104 on the gear reduction 48, which may be generally as
disclosed in the above-mentioned U.S. Patent, the oil need not sit
in the oil tank for ten seconds in order to de-aerate. Thus, the
size of the tank 90 may be made much smaller.
[0054] Applicant has discovered that oil is de-aerated by the
baffles 100 and gutter system and that a dwell time in the oil tank
to remove air bubbles may be less than five seconds. More
preferably, it may be less than or equal to about 3.0 seconds. This
allows the use of oil tank 90 to be of a size roughly equivalent to
the size utilized in prior non-geared gas turbine engines.
[0055] As an example, an oil tank that holds 25 to 35 quarts of oil
may be utilized on a geared gas turbine engine with 15,000 to
35,000 lbs in rated thrust at take-off. Further, an oil tank may be
35 quarts to 50 quarts of oil for an engine with 35,000 to 100,000
lbs in rated thrust at take-off.
[0056] Although an embodiment of this invention has been disclosed,
a worker of ordinary skill in this art would recognize that certain
modifications would come within the scope of this invention. For
that reason, the following claims should be studied to determine
the true scope and content of this invention.
* * * * *