CA1260277A - High mach number unducted fan engine - Google Patents
High mach number unducted fan engineInfo
- Publication number
- CA1260277A CA1260277A CA000507565A CA507565A CA1260277A CA 1260277 A CA1260277 A CA 1260277A CA 000507565 A CA000507565 A CA 000507565A CA 507565 A CA507565 A CA 507565A CA 1260277 A CA1260277 A CA 1260277A
- Authority
- CA
- Canada
- Prior art keywords
- engine
- excess
- pressure ratio
- ratio
- thrust
- Prior art date
- Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
- Expired
Links
- 239000000446 fuel Substances 0.000 claims description 8
- 239000007789 gas Substances 0.000 description 15
- 230000001141 propulsive effect Effects 0.000 description 9
- 238000013459 approach Methods 0.000 description 5
- 230000007423 decrease Effects 0.000 description 5
- 230000000694 effects Effects 0.000 description 4
- 239000000470 constituent Substances 0.000 description 2
- 238000012986 modification Methods 0.000 description 1
- 230000004048 modification Effects 0.000 description 1
Classifications
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
- F02C—GAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS
- F02C6/00—Plural gas-turbine plants; Combinations of gas-turbine plants with other apparatus; Adaptations of gas-turbine plants for special use
- F02C6/20—Adaptations of gas-turbine plants for driving vehicles
- F02C6/206—Adaptations of gas-turbine plants for driving vehicles the vehicles being airscrew driven
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
- F02K—JET-PROPULSION PLANTS
- F02K3/00—Plants including a gas turbine driving a compressor or a ducted fan
- F02K3/02—Plants including a gas turbine driving a compressor or a ducted fan in which part of the working fluid by-passes the turbine and combustion chamber
- F02K3/04—Plants including a gas turbine driving a compressor or a ducted fan in which part of the working fluid by-passes the turbine and combustion chamber the plant including ducted fans, i.e. fans with high volume, low pressure outputs, for augmenting the jet thrust, e.g. of double-flow type
- F02K3/062—Plants including a gas turbine driving a compressor or a ducted fan in which part of the working fluid by-passes the turbine and combustion chamber the plant including ducted fans, i.e. fans with high volume, low pressure outputs, for augmenting the jet thrust, e.g. of double-flow type with aft fan
Landscapes
- Engineering & Computer Science (AREA)
- Chemical & Material Sciences (AREA)
- Combustion & Propulsion (AREA)
- Mechanical Engineering (AREA)
- General Engineering & Computer Science (AREA)
- Structures Of Non-Positive Displacement Pumps (AREA)
- Control Of Vehicle Engines Or Engines For Specific Uses (AREA)
- Transition And Organic Metals Composition Catalysts For Addition Polymerization (AREA)
- Separation By Low-Temperature Treatments (AREA)
- Engine Equipment That Uses Special Cycles (AREA)
Abstract
HIGH MACH NUMBER UNDUCTED FAN ENGINE
ABSTRACT OF THE DISCLOSURE
An aircraft gas turbine engine capable of cruise operation at altitudes in excess of 30,000 feet and at mach numbers in excess of 0.6 is disclosed. The engine comprises a core engine including a compressor which is effective for producing a maximum pressure ratio in excess of 40:1. The engine further comprises a propulsor having an air bypass ratio between 10:1 and 60:1. The engine further comprises means for varying the pressure ratio so that the increase in pressure ratio from take-off to cruise is in excess of 20%.
ABSTRACT OF THE DISCLOSURE
An aircraft gas turbine engine capable of cruise operation at altitudes in excess of 30,000 feet and at mach numbers in excess of 0.6 is disclosed. The engine comprises a core engine including a compressor which is effective for producing a maximum pressure ratio in excess of 40:1. The engine further comprises a propulsor having an air bypass ratio between 10:1 and 60:1. The engine further comprises means for varying the pressure ratio so that the increase in pressure ratio from take-off to cruise is in excess of 20%.
Description
~6~
- l - 13DV-8753 HIGH MACH NUMBER UNDUCTED FAN ENGIN~
This invention relates generally to aircraft gas turbine engines and, more particularly, to turboprop engines capable of achieving a relatively high altitude and mach number.
BACKGROUND OF TH~ INVENTION
. .
Gas ~urbine engines for use on aircraft fall generally into three categories. These include turbojet, turbofan, and turboprop engines. All three engines have a core which includes a compressor, combustor, and turbine in serial flow relationship. The compressor compresses air entering the core to a relatively high pressure. This high pressure air is then mixed with fuel in the combustor and burned to form a high energy gas stream. This gas stream passes through the turbine where work is extracted to drive the compressor.
These three gas turbine engines differ in the manner in which they produce thrust. A turbojet uses the reactive force of the gas stream itself to provide the required thrust. A turbojet engine is capable of operating over a wide range of flight speeds including supersonic (Mach Number ~ 1). However, it is also characterized by relatively low thrust at take-off and generally poor fuel efficiency.
'~
~2~ ~77 13DV-~753 In contrast to the turbojet, the turbofan and turboprop engines generate thrust primarily through a propulsor radially disposed with respect to the core engine. In such engines, the high energy gas stream leaving the core turbine is expanded through a second turbine, known as a power turbine or low pressure turbine, which drives the fan or propulsor. Although some thrust is produced by the residual gas stream exiting the core nozzle, most of the thrust produced is generated by the propulsor. The ratio of the mass of air which passes through the propulsor (the air bypassing the core engine) to the mass of air passing through the core engine is known as the bypass ratio (~). The bypass ratio (~) is a rough measure of the ratio of thrust of the propulsor to thrust of the core exhaust.
Although the turboprop and turbofan engines operate under similar thermodynamic cycles, the two engines represent significantly different approaches to the design of gas turbine engines for aircraft. The conventional turboprop engine includes a relatively large diameter propeller with several blades for moving a relatively large volume of air and imparting to it a relatively small pressure rise. In contrast, the turbofan engine includes a significantly smaller diameter fan ~5 section encased within a cowling. The fan has a relatively large number of blades and imparts a relatively higher pressure rise to the volume of air passing therethrough, than the turboprop. For example, the pressure rise or pressure ratio of a typical propeller is on the order of 1.1 whereas the pressure ratio of a typical fan is about 1.7.
The differences between the three engines described above may be measured by the performance of each engine over a range of flight conditions, such as speed and altitude. Important measures of the performance of Z~7 the engine are engine efficiency and thrust. Engine efficiency includes a constituent representing how efficiently the heat energy of the fuel is converted into kinetic energy and a constituent measuring how efficiently the kinetic energy is converted to propulsive work. In other words, engine efficiency is made up of the thermal or internal efficiency of the engine and by propulsive or e~ternal efficiency of the engine.
Simply stated, propulsive efficiency is the ratio of the wor~ done by the engine over the useful energy imparted to the engine airflow. Algebraically, this may be expressed by the formula:
Ne = 2Vo/(VO + Vj), (1) where Ne = propulsive efficiency~
VO = aircraft speed, and Vj = exhaust velocity.
As can be seen, propulsive efficiency approaches 100% as Vj approaches VO That is, propulsive efficiency becomes high when the velocity of the exhaust gases approaches the velocity of the aircraft.
The thrust generated by an aircraft engine is proportional to the mass of air moved by the engine multiplied by the difference between the exit velocity and aircraft speed. This may be represented by:
Fn = Wa(Vj - vo)/gc~ (2) ` where Fn = net thrust, Wa = mass flow, and gc = a constant.
It is clear that large thrusts are obtainable by having a large mass flow or a large difference between the exhaust gas velocity and aircraft speed. Referring now to both equations (1) and (2), it becomes clear that for a given thrust it is more efficient to give a large mass of air a ~, small increase in velocity relative to the aircraft speed. Such is a characteristic of the turbofan and turboprop engine which distinguishes them from turbojet engines.
A difference between a conventional turboprop and turbofan engine is that the turboprop exhibits a rapid fall off in engine efficiency at higher aircraft speeds.
This is due to the supersonic flow relative to each propeller blade which increases the drag as the tip speed approaches the speed of sound. In contrast, the turbofan is capable of achie~ing higher overall efficiencies at high flight mach numbers because a diffuser section of the turbofan cowling reduces the speed of the incoming air below that of the aircraft speed. However, a limitation of the turbofan engine is that increased bypass ratios for increased propulsive efficiency require larger cowlings which result in excess weight and drag which seriously degrades aircraft fuel burn efficiency.
A recent improvement over the turbofan and turboprop engines is the so-called unducted fan engine, such as disclosed in Cdn. Serial Number 438,676 filed Oct. 7, 1983 - Johnson.- The unducted fan engine includes features of a conventional turboprop such as no cowling and variable pitch blades which are thin and swept to yield good efficiency at high aircraft speed as well as features of a turbofan such as increased number of blades per row and a lower tip diameter than a turboprop. In terms of bypass ratio, this places the unducted fan engine somewhere between a conventional turboprop and turbofan engine. For example, bypass ratios for the unducted fan on the order of 35:1 but in the range of 10:1 to 60:1 may be typical.
A complicating factor in the design of high bypass ratio engines, and particularly the conventional turboprop engines, is the phenomenon ~nown as lapse rate.
~60~77 13DV-8753 Lapse rate refers to the decrease in net thrust which occurs as the engine increases in Mach number and altitude. Referring again to equation (2), it can be seen that at take-off when VO is low that the net thrust will be the product of mass flow Wa and Vj. As the aircraft climbs in altitude, several things occur. First, VO increases with the speed of the aircraft with a smaller increase in Vj. Thus, the difference Vj -VO decreases. Second, with increased air speed the ram effect of air being pushed into the engine increases the density and mass flow of the air. However, with increased altitude the decrease in air density more than offsets the ram effect. Thus, mass flow Wa decreases as the cruise condition is reached.
High bypass ratio engines have a relatively large lapse rate. This means that a conventional high bypass ratio engine sized for take-off will be unable to achieve the mach number and altitude of an equivalently sized lower bypass ratio engine with less lapse. Although it is ~O possible within a certain range to design a high bypass ratio engine which meets a given flight speed and altitude, such an engine will be oversized for take-off.
In the past, turboprop engines accepted the large lapse rate and settled for the lower speeds and altitudes ~5 which could efficiently be attained. The combination of high propulsive e~ficiency, high bypass ratio, and a lapse rate which allows an engine to be sized for take-off conditions as well as meet thrust requirements for mach numbers in excess of 0.6 and altitudes in excess of 30,000 feet is now possible in connection with the unducted fan concept.
OBJ~CTS OF THE INVENTION
It is an object of the present invention to provide a new and improved aircraft gas turbine engine.
~ 7 ~ 13DV-8753 It is another object of the present invention to provide a new and improved unducted fan engine capable of achieving mach numbers in excess of 0.6 and altitudes in excess of 30,000 feet.
It is yet another object of the present invention to provide a relatively high bypass ratio engine with a reduced lapse rate.
SUMMARY OF THE INVENTION
The present invention is an aircraft gas turbine engine capable of cruise operation at altitudes in excess of 30,000 feet and at mach numbers in excess of 0.6. The engine comprises a core engine including a compressor which is effective for producing a maximum pressure ratio in excess of 40:1. The engine further comprises a propulsor having an air bypass ratio between 10:1 and 60:1. The engine further comprises means for varying the pressure ratio so that the increase in pressure ratio from take-off to cruise is in excess of 20%.
BRIEF DESCRIPTION OF THE DRAWINGS
FIGURE 1 is a graph showing the lapse rate of engines with different bypass ratios.
FIGURE 2 is a graph showing the effects of turbine inlet temperature on net thrust.
FIGURE 3 is a schematic view of a gas turbine engine according to the present invention.
DETAILED DESCRIPTION OF THE INVENTION
Figure 1 shows the lapse rate of engines with different bypass ratios (~). Lapse rate is represented as a plot of corrected thrust versus mach number. Corrected lZ~
thrust is represented by the Fn/Po, where Fn = net thrust and PO = ambient pressure. As can be seen, for high bypass ratios, the lapse rate or rate of decrease of corrected thrust for increasing mach numbers increases.
For B = 0 ~a pure turbojet) the lapse rate is relatively small compared to B = 5 (a typical turbofan) and B = 100 (a typical turboprop). The present invention relates to an unducted fan engine with B between 10:1 and 60:1 with a preferred range of between 20:1 and 40:1 and a preferred value within that range of approximately 35:1.
According to the present invention, the lapse rate curve of a typical turbofan represents the ideal case to match properly with efficient high speed aircraft design thrust requirements; with Fto representing the thrust required for take-off and FCr representing the thrust required for cruise at M = 0.8. Point C shows FCr for M = 0.8. It can be seen that a typical turbofan with a bypass ratio of about 5 provides a lapse rate curve such that Fto and FCr are in correct proportion to each other. An unducted fan engine designed with a core engine for providing Fto equivalent to that of the turbofan would follow a lapse rate curve shown by dashed line A in Figure 1. Properly sized for take-off, this engine would be unable to meet the thrust requirements FCr at M = 0.8 (shown by point B). In order to reduce the lapse rate curve A so that point B corresponds with point C, in accordance with this invention, it is proposed to modify the core engine.
Figure 2 shows the effect of changes in the corrected turbine inlet temperature on corrected net thrust. Figure 3 is a schematic view of a gas turbine engine according to one form of the present invention with specific engine stations called out. For reference, core engine 10 includes a compressor 12, combustor 1~, and turbine 16 in serial flow relationship. Located aft of ~2~ 77 turbine 16 is a power turbine 18 which is effective for turning unducted fan 20. A more detailed view of such an engine is shown in Cana~ian application Serial Number 438,676.
Station 2 is located just forward of compressor 12, station 3 is located at compressor discharge aft of compressor 12, and station 41 is located aft of combustor 14 and forward of-turbine 16. For a fixed al~itude and constant mach number, the temperature at station 2 or T2 is constant. T41 is a function of fuel flow to combustor 14. As shown by Figure 2, an increase in T
for a fixed T2 will result in an increase in corrected net thrust Fn/Po.
Ideally, in order to increase corrected net thrust, T41/T2 will be increased. However, increases in this ratio increase the speed of turbine 16 which increases the speed (Nc) of compressor 12. Increases in Nc result in an increase in the overall pressure ratio (P3/P2) of compressor 12, within the limits of the compressor design. Increasing Nc beyond the design point will result in a choke condition with no further increase in pressure ratio being realized. Pressure ratio increases result in temperature increases (T3) at compressor discharge. In the past, compressors have been designed for maximum pressure ratios of perhaps 34:1.
According to the present invention, it is proposed to provide a compressor with maximum pressure ratio in excess of 40:1. Preferably, the compressor will be operable over a range of between 6:1 to at least 45:1.
Take-of thrust will be achieved with a pressure ratio of approximately 30:1 and high altitude thrust will be obtained with an increase in overall pressure ratio to about 40:1 or higher. In a preferred embodiment, the increase in pressure ratio from take-off to cruise is in ~Z6VZ7 7 13DV-8753 g _ excess of 20%. Means for varying this pressure ratio are further provided in the form of a fuel flow control, i.e.
varying T41/T2 In operation, the present engine will be capable 5 of achieving cruise operation at altitudes in excess of 30,000 feet and mach numbers in excess of 0.6. Its propulsive efficiency (due to increased bypass ratio) should exceed any prior art engine capable of attaining these flight conditions. Furthermore, the engine will be minimally sized for both take-off and cruise condition and, therefore, be capable of attaining relatively low specific fuel consumption, and low weight.
It will be clear to those skilled in the art that numerous modifications, variations, and full and partial equivalents can now be undertaken without departing from the invention as limited only by the spirit and scope of the appended claims.
- l - 13DV-8753 HIGH MACH NUMBER UNDUCTED FAN ENGIN~
This invention relates generally to aircraft gas turbine engines and, more particularly, to turboprop engines capable of achieving a relatively high altitude and mach number.
BACKGROUND OF TH~ INVENTION
. .
Gas ~urbine engines for use on aircraft fall generally into three categories. These include turbojet, turbofan, and turboprop engines. All three engines have a core which includes a compressor, combustor, and turbine in serial flow relationship. The compressor compresses air entering the core to a relatively high pressure. This high pressure air is then mixed with fuel in the combustor and burned to form a high energy gas stream. This gas stream passes through the turbine where work is extracted to drive the compressor.
These three gas turbine engines differ in the manner in which they produce thrust. A turbojet uses the reactive force of the gas stream itself to provide the required thrust. A turbojet engine is capable of operating over a wide range of flight speeds including supersonic (Mach Number ~ 1). However, it is also characterized by relatively low thrust at take-off and generally poor fuel efficiency.
'~
~2~ ~77 13DV-~753 In contrast to the turbojet, the turbofan and turboprop engines generate thrust primarily through a propulsor radially disposed with respect to the core engine. In such engines, the high energy gas stream leaving the core turbine is expanded through a second turbine, known as a power turbine or low pressure turbine, which drives the fan or propulsor. Although some thrust is produced by the residual gas stream exiting the core nozzle, most of the thrust produced is generated by the propulsor. The ratio of the mass of air which passes through the propulsor (the air bypassing the core engine) to the mass of air passing through the core engine is known as the bypass ratio (~). The bypass ratio (~) is a rough measure of the ratio of thrust of the propulsor to thrust of the core exhaust.
Although the turboprop and turbofan engines operate under similar thermodynamic cycles, the two engines represent significantly different approaches to the design of gas turbine engines for aircraft. The conventional turboprop engine includes a relatively large diameter propeller with several blades for moving a relatively large volume of air and imparting to it a relatively small pressure rise. In contrast, the turbofan engine includes a significantly smaller diameter fan ~5 section encased within a cowling. The fan has a relatively large number of blades and imparts a relatively higher pressure rise to the volume of air passing therethrough, than the turboprop. For example, the pressure rise or pressure ratio of a typical propeller is on the order of 1.1 whereas the pressure ratio of a typical fan is about 1.7.
The differences between the three engines described above may be measured by the performance of each engine over a range of flight conditions, such as speed and altitude. Important measures of the performance of Z~7 the engine are engine efficiency and thrust. Engine efficiency includes a constituent representing how efficiently the heat energy of the fuel is converted into kinetic energy and a constituent measuring how efficiently the kinetic energy is converted to propulsive work. In other words, engine efficiency is made up of the thermal or internal efficiency of the engine and by propulsive or e~ternal efficiency of the engine.
Simply stated, propulsive efficiency is the ratio of the wor~ done by the engine over the useful energy imparted to the engine airflow. Algebraically, this may be expressed by the formula:
Ne = 2Vo/(VO + Vj), (1) where Ne = propulsive efficiency~
VO = aircraft speed, and Vj = exhaust velocity.
As can be seen, propulsive efficiency approaches 100% as Vj approaches VO That is, propulsive efficiency becomes high when the velocity of the exhaust gases approaches the velocity of the aircraft.
The thrust generated by an aircraft engine is proportional to the mass of air moved by the engine multiplied by the difference between the exit velocity and aircraft speed. This may be represented by:
Fn = Wa(Vj - vo)/gc~ (2) ` where Fn = net thrust, Wa = mass flow, and gc = a constant.
It is clear that large thrusts are obtainable by having a large mass flow or a large difference between the exhaust gas velocity and aircraft speed. Referring now to both equations (1) and (2), it becomes clear that for a given thrust it is more efficient to give a large mass of air a ~, small increase in velocity relative to the aircraft speed. Such is a characteristic of the turbofan and turboprop engine which distinguishes them from turbojet engines.
A difference between a conventional turboprop and turbofan engine is that the turboprop exhibits a rapid fall off in engine efficiency at higher aircraft speeds.
This is due to the supersonic flow relative to each propeller blade which increases the drag as the tip speed approaches the speed of sound. In contrast, the turbofan is capable of achie~ing higher overall efficiencies at high flight mach numbers because a diffuser section of the turbofan cowling reduces the speed of the incoming air below that of the aircraft speed. However, a limitation of the turbofan engine is that increased bypass ratios for increased propulsive efficiency require larger cowlings which result in excess weight and drag which seriously degrades aircraft fuel burn efficiency.
A recent improvement over the turbofan and turboprop engines is the so-called unducted fan engine, such as disclosed in Cdn. Serial Number 438,676 filed Oct. 7, 1983 - Johnson.- The unducted fan engine includes features of a conventional turboprop such as no cowling and variable pitch blades which are thin and swept to yield good efficiency at high aircraft speed as well as features of a turbofan such as increased number of blades per row and a lower tip diameter than a turboprop. In terms of bypass ratio, this places the unducted fan engine somewhere between a conventional turboprop and turbofan engine. For example, bypass ratios for the unducted fan on the order of 35:1 but in the range of 10:1 to 60:1 may be typical.
A complicating factor in the design of high bypass ratio engines, and particularly the conventional turboprop engines, is the phenomenon ~nown as lapse rate.
~60~77 13DV-8753 Lapse rate refers to the decrease in net thrust which occurs as the engine increases in Mach number and altitude. Referring again to equation (2), it can be seen that at take-off when VO is low that the net thrust will be the product of mass flow Wa and Vj. As the aircraft climbs in altitude, several things occur. First, VO increases with the speed of the aircraft with a smaller increase in Vj. Thus, the difference Vj -VO decreases. Second, with increased air speed the ram effect of air being pushed into the engine increases the density and mass flow of the air. However, with increased altitude the decrease in air density more than offsets the ram effect. Thus, mass flow Wa decreases as the cruise condition is reached.
High bypass ratio engines have a relatively large lapse rate. This means that a conventional high bypass ratio engine sized for take-off will be unable to achieve the mach number and altitude of an equivalently sized lower bypass ratio engine with less lapse. Although it is ~O possible within a certain range to design a high bypass ratio engine which meets a given flight speed and altitude, such an engine will be oversized for take-off.
In the past, turboprop engines accepted the large lapse rate and settled for the lower speeds and altitudes ~5 which could efficiently be attained. The combination of high propulsive e~ficiency, high bypass ratio, and a lapse rate which allows an engine to be sized for take-off conditions as well as meet thrust requirements for mach numbers in excess of 0.6 and altitudes in excess of 30,000 feet is now possible in connection with the unducted fan concept.
OBJ~CTS OF THE INVENTION
It is an object of the present invention to provide a new and improved aircraft gas turbine engine.
~ 7 ~ 13DV-8753 It is another object of the present invention to provide a new and improved unducted fan engine capable of achieving mach numbers in excess of 0.6 and altitudes in excess of 30,000 feet.
It is yet another object of the present invention to provide a relatively high bypass ratio engine with a reduced lapse rate.
SUMMARY OF THE INVENTION
The present invention is an aircraft gas turbine engine capable of cruise operation at altitudes in excess of 30,000 feet and at mach numbers in excess of 0.6. The engine comprises a core engine including a compressor which is effective for producing a maximum pressure ratio in excess of 40:1. The engine further comprises a propulsor having an air bypass ratio between 10:1 and 60:1. The engine further comprises means for varying the pressure ratio so that the increase in pressure ratio from take-off to cruise is in excess of 20%.
BRIEF DESCRIPTION OF THE DRAWINGS
FIGURE 1 is a graph showing the lapse rate of engines with different bypass ratios.
FIGURE 2 is a graph showing the effects of turbine inlet temperature on net thrust.
FIGURE 3 is a schematic view of a gas turbine engine according to the present invention.
DETAILED DESCRIPTION OF THE INVENTION
Figure 1 shows the lapse rate of engines with different bypass ratios (~). Lapse rate is represented as a plot of corrected thrust versus mach number. Corrected lZ~
thrust is represented by the Fn/Po, where Fn = net thrust and PO = ambient pressure. As can be seen, for high bypass ratios, the lapse rate or rate of decrease of corrected thrust for increasing mach numbers increases.
For B = 0 ~a pure turbojet) the lapse rate is relatively small compared to B = 5 (a typical turbofan) and B = 100 (a typical turboprop). The present invention relates to an unducted fan engine with B between 10:1 and 60:1 with a preferred range of between 20:1 and 40:1 and a preferred value within that range of approximately 35:1.
According to the present invention, the lapse rate curve of a typical turbofan represents the ideal case to match properly with efficient high speed aircraft design thrust requirements; with Fto representing the thrust required for take-off and FCr representing the thrust required for cruise at M = 0.8. Point C shows FCr for M = 0.8. It can be seen that a typical turbofan with a bypass ratio of about 5 provides a lapse rate curve such that Fto and FCr are in correct proportion to each other. An unducted fan engine designed with a core engine for providing Fto equivalent to that of the turbofan would follow a lapse rate curve shown by dashed line A in Figure 1. Properly sized for take-off, this engine would be unable to meet the thrust requirements FCr at M = 0.8 (shown by point B). In order to reduce the lapse rate curve A so that point B corresponds with point C, in accordance with this invention, it is proposed to modify the core engine.
Figure 2 shows the effect of changes in the corrected turbine inlet temperature on corrected net thrust. Figure 3 is a schematic view of a gas turbine engine according to one form of the present invention with specific engine stations called out. For reference, core engine 10 includes a compressor 12, combustor 1~, and turbine 16 in serial flow relationship. Located aft of ~2~ 77 turbine 16 is a power turbine 18 which is effective for turning unducted fan 20. A more detailed view of such an engine is shown in Cana~ian application Serial Number 438,676.
Station 2 is located just forward of compressor 12, station 3 is located at compressor discharge aft of compressor 12, and station 41 is located aft of combustor 14 and forward of-turbine 16. For a fixed al~itude and constant mach number, the temperature at station 2 or T2 is constant. T41 is a function of fuel flow to combustor 14. As shown by Figure 2, an increase in T
for a fixed T2 will result in an increase in corrected net thrust Fn/Po.
Ideally, in order to increase corrected net thrust, T41/T2 will be increased. However, increases in this ratio increase the speed of turbine 16 which increases the speed (Nc) of compressor 12. Increases in Nc result in an increase in the overall pressure ratio (P3/P2) of compressor 12, within the limits of the compressor design. Increasing Nc beyond the design point will result in a choke condition with no further increase in pressure ratio being realized. Pressure ratio increases result in temperature increases (T3) at compressor discharge. In the past, compressors have been designed for maximum pressure ratios of perhaps 34:1.
According to the present invention, it is proposed to provide a compressor with maximum pressure ratio in excess of 40:1. Preferably, the compressor will be operable over a range of between 6:1 to at least 45:1.
Take-of thrust will be achieved with a pressure ratio of approximately 30:1 and high altitude thrust will be obtained with an increase in overall pressure ratio to about 40:1 or higher. In a preferred embodiment, the increase in pressure ratio from take-off to cruise is in ~Z6VZ7 7 13DV-8753 g _ excess of 20%. Means for varying this pressure ratio are further provided in the form of a fuel flow control, i.e.
varying T41/T2 In operation, the present engine will be capable 5 of achieving cruise operation at altitudes in excess of 30,000 feet and mach numbers in excess of 0.6. Its propulsive efficiency (due to increased bypass ratio) should exceed any prior art engine capable of attaining these flight conditions. Furthermore, the engine will be minimally sized for both take-off and cruise condition and, therefore, be capable of attaining relatively low specific fuel consumption, and low weight.
It will be clear to those skilled in the art that numerous modifications, variations, and full and partial equivalents can now be undertaken without departing from the invention as limited only by the spirit and scope of the appended claims.
Claims (3)
1. An aircraft gas turbine engine capable of cruise operation at altitudes in excess of 30,000 feet and at Mach numbers in excess of 0.6 comprising:
a core engine including a compressor effective for producing a maximum pressure ratio in excess of 40:1;
a propulsor having an air bypass ratio between 10:1 and 60:1; and a fuel flow control means effective for varying said pressure ratio so that the increase in pressure ratio from take-off to cruise is in excess of 20%.
a core engine including a compressor effective for producing a maximum pressure ratio in excess of 40:1;
a propulsor having an air bypass ratio between 10:1 and 60:1; and a fuel flow control means effective for varying said pressure ratio so that the increase in pressure ratio from take-off to cruise is in excess of 20%.
2. An engine, as recited in claim 1, wherein said propulsor has an air bypass ratio between 20:1 and 40:1.
3. An engine, as recited in claim 1, capable of cruise operation at an altitude of 35,000 feet and a Mach number of 0.8, said propulsor having an air bypass ratio approximately equal to 35:1 and said compressor having a maximum pressure ratio of approximately 45:1.
Applications Claiming Priority (2)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| US73002385A | 1985-05-03 | 1985-05-03 | |
| US730,023 | 1985-05-03 |
Publications (1)
| Publication Number | Publication Date |
|---|---|
| CA1260277A true CA1260277A (en) | 1989-09-26 |
Family
ID=24933598
Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| CA000507565A Expired CA1260277A (en) | 1985-05-03 | 1986-04-25 | High mach number unducted fan engine |
Country Status (7)
| Country | Link |
|---|---|
| JP (1) | JPS61283755A (en) |
| CA (1) | CA1260277A (en) |
| DE (1) | DE3615006A1 (en) |
| FR (1) | FR2581425A1 (en) |
| GB (1) | GB2174761B (en) |
| IT (1) | IT1191772B (en) |
| SE (1) | SE8602031L (en) |
Families Citing this family (3)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US20070260424A1 (en) * | 2006-05-05 | 2007-11-08 | Harold Brown | Methods and apparatus for estimating engine thrust |
| US8082727B2 (en) | 2008-02-26 | 2011-12-27 | United Technologies Corporation | Rear propulsor for a variable cycle gas turbine engine |
| US8127528B2 (en) | 2008-02-26 | 2012-03-06 | United Technologies Corporation | Auxiliary propulsor for a variable cycle gas turbine engine |
Family Cites Families (2)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| GB1487324A (en) * | 1973-11-15 | 1977-09-28 | Rolls Royce | Gas turbine engines |
| US4502275A (en) * | 1983-10-13 | 1985-03-05 | Avco Corporation | Compressor air bleed override control system |
-
1986
- 1986-04-22 GB GB8609742A patent/GB2174761B/en not_active Expired
- 1986-04-24 FR FR8605936A patent/FR2581425A1/en not_active Withdrawn
- 1986-04-25 CA CA000507565A patent/CA1260277A/en not_active Expired
- 1986-04-28 JP JP61097044A patent/JPS61283755A/en active Pending
- 1986-04-30 IT IT20278/86A patent/IT1191772B/en active
- 1986-04-30 SE SE8602031A patent/SE8602031L/en not_active Application Discontinuation
- 1986-05-02 DE DE19863615006 patent/DE3615006A1/en not_active Withdrawn
Also Published As
| Publication number | Publication date |
|---|---|
| IT1191772B (en) | 1988-03-23 |
| IT8620278A1 (en) | 1987-10-30 |
| SE8602031L (en) | 1986-11-04 |
| SE8602031D0 (en) | 1986-04-30 |
| GB2174761B (en) | 1989-09-06 |
| FR2581425A1 (en) | 1986-11-07 |
| DE3615006A1 (en) | 1986-11-13 |
| IT8620278A0 (en) | 1986-04-30 |
| GB8609742D0 (en) | 1986-05-29 |
| GB2174761A (en) | 1986-11-12 |
| JPS61283755A (en) | 1986-12-13 |
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| MKEX | Expiry |