Patent Publication Number: US-7216475-B2

Title: Aft FLADE engine

Description:
BACKGROUND OF THE INVENTION 
   1. Field of the Invention 
   This invention relates to aircraft gas turbine engines and, more particularly, to FLADE engines. 
   2. Description of Related Art 
   High performance variable cycle gas turbine engines are being designed because of their unique ability to operate efficiently at various thrust settings and flight speeds both subsonic and supersonic. An important feature of the variable cycle gas turbine engine which contributes to its high performance is its capability of maintaining a substantially constant inlet airflow as its thrust is varied. This feature leads to important performance advantages under less than full power engine settings or maximum thrust conditions, such as during subsonic cruise. 
   One particular type of variable cycle engine called a FLADE engine (FLADE being an acronym for “fan on blade”) is characterized by an outer fan driven by a radially inner fan and discharging its FLADE air into an outer fan duct which is generally co-annular with and circumscribes an inner fan duct circumscribing the inner fan. One such engine, disclosed in U.S. Pat. No. 4,043,121, entitled “Two Spool Variable Cycle Engine”, by Thomas et al., provides a FLADE fan and outer fan duct within which variable guide vanes control the cycle variability by controlling the amount of air passing through the FLADE outer fan duct. Other high performance aircraft variable cycle gas turbine FLADE engines capable of maintaining an essentially constant inlet airflow over a relatively wide range of thrust at a given set of subsonic flight ambient conditions such as altitude and flight Mach No. in order to avoid spillage drag and to do so over a range of flight conditions have been studied. This capability is particularly needed for subsonic part power engine operating conditions. Examples of these are disclosed in U.S. Pat. No. 5,404,713, entitled “Spillage Drag and Infrared Reducing Flade Engine”, U.S. Pat. No. 5,402,963, entitled “Acoustically Shielded Exhaust System for High Thrust Jet Engines”, U.S. Pat. No. 5,261,227, entitled “Variable Specific Thrust Turbofan Engine”, and European Patent No. EP0,567,277, entitled “Bypass Injector Valve For Variable Cycle Aircraft Engines”. A FLADE aircraft gas turbine engine with counter-rotatable fans is disclosed in U.S. patent application Ser. No. (10/647,881), entitled “FLADE GAS TURBINE ENGINE WITH COUNTER-ROTATABLE FANS”. 
   FLADE engines have the fan blade attached to one of the front fans. This can lead to low pressure spool designs that are compromised because of the limitations in rotor speeds and increased stresses caused by the FLADE blade attachment and location. The front fan mounted FLADE fan blades also are difficult to adapt to present engines or engine designs. It would be very expensive to adapt an existing engine to test a front fan mounted FLADE fan. It would be difficult to demonstrate some of the system benefits offered by a FLADE engine concept at a reasonable cost relative to that of a new low pressure system or defined around an existing core engine. 
   It is highly desirable to have a FLADE engine that allows a low pressure spool design that is uncompromised because of limitations in rotor speeds and increased stresses caused by the FLADE blade attachment and location. It is highly desirable to have an engine in which FLADE fan blades are not difficult to adapt to present engines or engine designs and that would not be very expensive to adapt to an existing engine to test as compared to a front fan mounted FLADE fan. It is also desirable to be able to demonstrate some of the system benefits offered by a FLADE engine concept without a great deal of difficulty and at a reasonable cost relative to that of a new low pressure system or defined around an existing core engine. 
   SUMMARY OF THE INVENTION 
   An aft FLADE gas turbine engine includes a fan section drivenly connected to a low pressure turbine section, a core engine located between the fan section and the low pressure turbine section, a fan bypass duct circumscribing the core engine and in fluid communication with the fan section, a mixer downstream of the low pressure turbine section and in fluid communication with the fan bypass duct, and an aft FLADE turbine downstream of the mixer. At least one row of aft FLADE fan blades is disposed radially outwardly of and drivenly connected to the aft FLADE turbine. The row of FLADE fan blades radially extend across a FLADE duct circumscribing the fan section. 
   More particular embodiments of the engine include a row of variable first FLADE vanes radially extending across the FLADE duct axially forwardly of the row of FLADE fan blades. One embodiment of the engine further includes a fan inlet to the fan section and an annular FLADE inlet to the FLADE duct arranged so that the FLADE inlet is axially located substantially aftwardly of the fan section and, in a more particular embodiment, the FLADE inlet is axially located aftwardly of the core engine. The aft FLADE turbine may be connected to and rotatable with a low pressure turbine of the low pressure turbine section or may be a free turbine. The engine may incorporate a variable area turbine nozzle with variable turbine nozzle vanes located aft and downstream of the mixer and the low pressure turbine. 
   A power extraction apparatus may be placed within the engine and drivenly connected to the aft FLADE turbine. In one embodiment, the power extraction apparatus may be located in a hollow engine nozzle centerbody of the engine located aft and downstream of the aft FLADE turbine. One embodiment of the power extraction apparatus is an electrical generator drivenly connected through a speed increasing gearbox to the aft FLADE turbine. Another embodiment of the power extraction apparatus is a power takeoff assembly including a housing disposed within the engine such as in the hollow engine nozzle centerbody and having a power takeoff shaft drivenly connected to the aft FLADE turbine through a right angle gearbox. 
   A variable or fixed throat area engine nozzle may be incorporated downstream and axially aft of the mixer and the fan bypass duct. Another more particular embodiment of the engine includes a plurality of circumferentially disposed hollow struts in fluid flow communication with the FLADE duct and a substantially hollow centerbody supported by and in fluid flow communication with the hollow struts. Cooling apertures in the centerbody and in a wall of the engine nozzle downstream of the variable throat area are in fluid communication with the FLADE duct. 
   A variable area FLADE air nozzle including an axially translatable plug within the hollow centerbody and a radially outwardly positioned fixed nozzle cowling of the centerbody may also be used in the engine. Aft thrust augmenting afterburners may be incorporated aft and downstream of the aft FLADE turbine. A forward afterburner may be axially disposed between the mixer and the aft FLADE turbine to provide additional energy upon demand to the aft FLADE turbine and additional power to the row of aft FLADE fan blades and the power extraction apparatus such as the electrical generator or the power takeoff assembly. 
   The aft FLADE gas turbine engine may be used within a fuselage of the aircraft. FLADE air intakes and an engine air intake may be mounted flush with respect to the fuselage. The FLADE air intakes are axially offset from the engine air intake. The engine air intake may be connected to and in fluid communication with the fan inlet by an engine fixed inlet duct. The FLADE air intakes may be connected to and in fluid communication with the FLADE inlets by FLADE fixed inlet ducts. Inlet duct passages of the engine and the FLADE fixed inlet ducts respectively may be two-dimensional and terminating in transition sections between the inlet duct passages and the fan and FLADE inlets respectively. 
   The aft FLADE turbine allows a FLADE engine to have a low pressure spool design that is uncompromised because of limitations in rotor speeds and increased stresses caused by the FLADE blade attachment and location. FLADE fan blades mounted on the aft FLADE turbine are not difficult to adapt to present engines or engine designs. The aft FLADE turbine is not very expensive to adapt to an existing engine to test as compared to a front fan mounted FLADE fan. It is also easier to demonstrate some of the system benefits offered by a FLADE engine concept without a great deal of difficulty and at a reasonable cost relative to that of a new low pressure system or defined around an existing core engine. 

   
     BRIEF DESCRIPTION OF THE DRAWINGS 
     The foregoing aspects and other features of the invention are explained in the following description, taken in connection with the accompanying drawings where: 
       FIG. 1  is a schematical cross-sectional view illustration of a FLADE aircraft gas turbine engine with a single direction of rotation fan section and an aft FLADE blade and turbine. 
       FIG. 2  is an alternative schematical cross-sectional view illustration of the engine in  FIG. 2  with exhaust nozzle cooling. 
       FIG. 3  is a schematical cross-sectional view illustration of an aircraft gas turbine engine with an aft FLADE blade and turbine and a short FLADE duct. 
       FIG. 4  is a schematical cross-sectional view illustration of the short aft duct FLADE gas turbine engine in  FIG. 3  installed in an aircraft. 
       FIG. 5  is a schematical cross-sectional view illustration of a short duct aft FLADE turbine module having an aft FLADE turbine drivingly connected to an internally mounted electrical generator. 
       FIG. 6  is a schematical cross-sectional view illustration of a FLADE aircraft gas turbine engine with a first afterburner upstream of a free aft FLADE turbine. 
       FIG. 7  is a schematical cross-sectional view illustration of an aft FLADE aircraft gas turbine engine with a variable area turbine nozzle and a thrust augmenting afterburner downstream of an aft FLADE turbine. 
       FIG. 8  is a schematical cross-sectional view illustration of a FLADE aircraft gas turbine engine with counter rotatable fans and an aft FLADE blade and turbine. 
       FIG. 9  is a schematical cross-sectional view illustration of a FLADE aircraft gas turbine engine with an aft FLADE blade and turbine driving connected to a power takeoff shaft. 
       FIG. 10  is a schematical cross-sectional view illustration of a FLADE aircraft gas turbine engine with an aft FLADE blade and turbine driving connected to an electrical generator located within the engine. 
       FIG. 11  is a schematical cross-sectional view illustration of a FLADE aircraft gas turbine engine with an aft FLADE blade and turbine driving connected to two electrical generators located within the engine. 
   

   DETAILED DESCRIPTION OF THE INVENTION 
   Schematically illustrated in cross-section in  FIG. 1  is an aircraft aft FLADE engine  1  having a fan section  115  with a single direction of rotation fan  330  downstream of variable inlet guide vanes  4 . Downstream and axially aft of the fan section  115  is a core engine  18  having an annular core engine inlet  17  and a generally axially extending axis or centerline  12  generally extending forward  14  and aft  16 . A fan bypass duct  40  located downstream and axially aft of the fan section  115  circumscribes the core engine  18 . A FLADE duct  3  circumscribes the fan section  115  and the fan bypass duct  40 . Fairings  190  disposed across the FLADE duct  3  surround variable vane shafts  194  passing through the FLADE duct  3  that are used to vary and control the pitch of the variable inlet guide vanes  4 . 
   The core engine  18  includes, in downstream serial axial flow relationship, a core driven fan  37  having a row of core driven fan blades  36 , a high pressure compressor  20 , a combustor  22 , and a high pressure turbine  23  having a row of high pressure turbine blades  24 . A high pressure shaft  26 , disposed coaxially about the centerline  12  of the engine  1 , fixedly interconnects the high pressure compressor  20  and the high pressure turbine blades  24 . The combination or assembly of the core driven fan  37  and the high pressure compressor  20  drivenly connected to the high pressure turbine  23  by the high pressure shaft  26  is designated a high pressure spool  47 . The core engine  18  is effective for generating combustion gases. Pressurized air from the high pressure compressor  20  is mixed with fuel in the combustor  22  and ignited, thereby, generating combustion gases. Some work is extracted from these gases by the high pressure turbine blades  24  which drives the core driven fan  37  and the high pressure compressor  20 . The high pressure shaft  26  rotates the core driven fan  37  having a single row of circumferentially spaced apart core driven fan blades  36  having generally radially outwardly located blade tip sections  38  separated from generally radially inwardly located blade hub sections  39  by an annular fan shroud  108 . 
   The combustion gases are discharged from the core engine  18  into single direction of rotation low pressure turbine  319  having at least one row of low pressure turbine blades  328 . The low pressure turbine  319  is drivingly connected to the single direction of rotation fan  330  by a low pressure shaft  321 , the combination or assembly being designated a first low pressure spool  240 . The single direction of rotation fan  330  has at least one row of generally radially outwardly extending and circumferentially spaced-apart fan blades  333 . 
   A mixer  49 , illustrated as a lobed or chute mixer, is disposed downstream of and at an aft end of the fan bypass duct  40  and downstream of and aft of the low pressure turbine blades  328  of the low pressure turbine  319 . The mixer  49  is used to mix bypass air  78  with core discharge air  70  from the low pressure turbine  319  to form a mixed flow  188 . One alternative version of the mixer  49  is an aft variable area bypass injector (VABI) door disposed at an aft end of the fan bypass duct  40  to mix bypass air  78  with core discharge air  70 . Aft VABI doors are typically circumferentially disposed between hollow struts  208  which structurally support and flow air to a hollow engine nozzle centerbody  72 . 
   Exhaust gases from the mixer  49  are directed through an aft FLADE turbine  160  having a plurality of FLADE turbine blades  254  and located downstream and aft of the mixer  49 . A FLADE fan  2  includes at least one row of aft FLADE fan blades  5  extend radially outwardly from and are drivenly connected to the aft FLADE turbine  160  across a FLADE duct  3  disposed radially outwardly of and circumscribing the fan bypass duct  40 . A FLADE airflow  80  is powered by the FLADE fan blades  5  and put to use downstream of the FLADE fan blades  5 . The FLADE fan blades  5  extend radially outwardly from an annular rotatable FLADE turbine shroud  250  attached to and circumscribing the FLADE turbine blades  254  of the aft FLADE turbine  160 . The FLADE turbine shroud  250  separates the FLADE fan blades  5  from the FLADE turbine blades  254 . 
   The embodiment of the engine  1  illustrated in  FIGS. 1 and 2  have what is referred to as a long FLADE duct  3  that extends forwardly from the FLADE fan blades  5  to a fan inlet  11  to the fan section  115 . The fan inlet  11  together with an annular FLADE inlet  8  to the FLADE duct  3  generally form an engine inlet  13 . The aft FLADE turbine  160  is illustrated, in  FIGS. 1 and 2 , drivingly connected to the low pressure shaft  321  and, thus, rotatable with the low pressure turbine  319 . Alternatively, the aft FLADE turbine  160  may be a free turbine not connected to either of the low or high pressure shafts as illustrated in  FIGS. 5–7  and  9 – 10 . 
   Referring to  FIG. 1 , a row of variable first FLADE vanes  6  is located within the FLADE duct  3  axially forwardly and upstream of the row of FLADE fan blades  5  to control a FLADE airflow  80  through the FLADE duct  3 . Variable vane shafts  194  are used to vary and control the pitch of the variable first FLADE vanes  6 . A row of second FLADE vanes  7  is illustrated in  FIG. 1  as being fixed but may be variable. The row of second FLADE vanes  7  is also located within the FLADE duct  3  but axially aftwardly and downstream of the row of FLADE fan blades  5 . The second FLADE vanes  7  are used to deswirl the FLADE airflow  80 . 
   Too much air or too little air passing through the fan inlet  11  is detrimental to aircraft system performance. The aft FLADE fan blades  5  of the FLADE fan  2  and the FLADE duct  3  are designed and operated to help manage the inlet airflow delivered by the inlet to the fans. The aft FLADE turbine  160  and the FLADE fan blades  5  of the FLADE fan  2  are used to allow the engine  1  to be operated in an optimum and efficient manner at all flight conditions. The fan inlet  11  is sized to receive essentially full engine airflow  15  of the engine at full power conditions. The FLADE airflow  80  controlled by the variable first FLADE vanes  6  can be scheduled from a maximum level to a minimum level at a constant rotor speed of the FLADE turbine  160 . 
   An optional variable area turbine nozzle  180  with variable turbine nozzle vanes  182  is illustrated in  FIG. 1  located downstream of the mixer  49  and the low pressure turbine  319  and immediately upstream of and at an entrance to the aft FLADE turbine  160 . Variable area nozzle vane shafts  192  that are used to vary and control the pitch of the variable turbine nozzle vanes  182  pass through the variable vane shafts  194  that are used to vary and control the pitch of the variable first FLADE vanes  6 . 
   Referring to  FIGS. 1 and 2 , a first bypass inlet  42  to the fan bypass duct  40  is disposed axially between the fan section  115  and the core driven fan  37 . The fan blades  333  of the fan  330  radially extend across a first fan duct  138 . A row of circumferentially spaced-apart fan stator vanes  35  radially extend across the first fan duct  138 , downstream of the fan blades  333 , and axially between the fan blades  333  and the first bypass inlet  42  to the fan bypass duct  40 . The row of the core driven fan blades  36  of the core driven fan  37  radially extend across an annular second fan duct  142 . The second fan duct  142  begins axially aft of the first bypass inlet  42  and is disposed radially inwardly of the fan bypass duct  40 . An annular first flow splitter  45  is radially disposed between the first bypass inlet  42  and the second fan duct  142 . 
   The full engine airflow  15  is split between the FLADE inlet  8  and the fan inlet  11 . A fan airflow  50  passes through the fan inlet  11  and then the fan section  115 . A first bypass air portion  52  of the fan airflow  50  passes through the first bypass inlet  42  of the fan bypass duct  40  when a front variable area bypass injector (VABI) door  44  in the first bypass inlet  42  is open and with the remaining air portion  54  passing through the core driven fan  37  and its row of core driven fan blades  36 . 
   A row of circumferentially spaced-apart core driven fan stator vanes  34  within the second fan duct  142  are disposed axially between the row of second fan blades  32  and the core driven fan blades  36  of the core driven fan  37 . The row of the core driven fan stator vanes  34  and the core driven fan blades  36  of the core driven fan  37  are radially disposed across the second fan duct  142 . A vane shroud  106  divides the core driven fan stator vanes  34  into radially inner and outer vane hub and tip sections  85  and  84 , respectively. The fan shroud  108  divides the core driven fan blades  36  into the radially inner and outer blade hub and tip sections  39  and  38 , respectively. 
   A second bypass airflow portion  56  is directed through a fan tip duct  146  across the vane tip sections  84  of the core driven fan stator vanes  34  and across the blade tip sections  38  of the core driven fan blades  36  into a second bypass inlet  46  of a second bypass duct  58  to the fan bypass duct  40 . An optional middle variable area bypass injector (VABI) door  83  may be disposed at an aft end of the second bypass duct  58  for modulating flow through the second bypass inlet  46  to the fan bypass duct  40 . 
   The fan tip duct  146  includes the vane and fan shrouds  106  and  108  and a second flow splitter  55  at a forward end of the vane shroud  106 . First and second varying means  91  and  92  are provided for independently varying flow areas of the vane hub and tip sections  85  and  84 , respectively. Exemplary first and second varying means  91  and  92  include independently variable vane hub and tip sections  85  and  84 , respectively (see U.S. Pat. No. 5,806,303). The independently variable vane hub and tip sections  85  and  84  designs may include having the entire vane hub and tip sections  85  and  84  be independently pivotable. Other possible designs are disclosed in U.S. Pat. Nos. 5,809,772 and 5,988,890. 
   Another embodiment of the independently variable vane hub and tip sections  85  and  84  includes pivotable trailing-edge hub and tip flaps  86  and  88  of the independently variable vane hub and tip sections  85  and  84 . The first and second varying means  91  and  92  can include independently pivoting flaps. Alternative varying means for non-pivotable, fan stator vane designs include axially moving unison rings and those means known for mechanical clearance control in jet engines (i.e., mechanically moving circumferentially surrounding shroud segments radially towards and away from a row of rotor blade tips to maintain a constant clearance despite different rates of thermal expansion and contraction). Additional such varying means for non-pivotable, fan stator vane designs include those known for extending and retracting wing flaps on airplanes and the like. 
   Exemplary first and second varying means  91  and  92 , illustrated in  FIG. 1 , include an inner shaft  94  coaxially disposed within an outer shaft  96 . The inner shaft  94  is rotated by a first lever arm  98  actuated by a first unison ring  100 . The outer shaft  96  is rotated by a second lever arm  102  actuated by a second unison ring  104 . The inner shaft  94  is attached to the pivotable trailing edge hub flap  86  of the vane hub section  85  of the fan stator vane  34 . The outer shaft  96  is attached to the pivotable trailing edge tip flap  88  of the vane tip section  84  of the fan stator vane  34 . It is noted that the lever arms  98  and  102  and the unison rings  100  and  104  are all disposed radially outward of the fan stator vanes  34 . Other such pivoting means include those known for pivoting variable stator vanes of high pressure compressors in jet engines and the like. 
   Referring to  FIGS. 1 and 2  by way of example, a variable throat area engine nozzle  218 , having a variable throat area A 8 , is downstream and axially aft of the aft FLADE turbine  160  and the fan bypass duct  40 . The engine nozzle  218  includes an axially translatable radially outer annular convergent and divergent wall  220  spaced radially outwardly apart from a radially fixed and axially translatable annular inner wall  222  on the centerbody  72 . The translatable radially outer annular convergent and divergent wall  220  controls a throat area A 8  between the outer annular convergent and divergent wall  220  and the radially fixed and axially translatable annular inner wall  222 . The translatable radially outer annular convergent and divergent wall  220  also controls a nozzle exit area A 9  of the engine nozzle  218 . Alternatively, a variable throat area convergent/divergent nozzle with flaps may be used as disclosed in U.S. Pat. No. 5,404,713. 
   The plurality of circumferentially disposed hollow struts  208  are in fluid communication with and operable to receive air from the FLADE duct  3 . The hollow struts  208  structurally support and flow air to the centerbody  72  which is substantially hollow. A variable area FLADE air nozzle  213  includes an axially translatable plug  172  which cooperates with a radially outwardly positioned fixed nozzle cowling  174  of the centerbody  72  to exhaust FLADE airflow  80  received from the hollow struts  208  and return work to the engine in the form of thrust. 
     FIGS. 2 ,  3 ,  6 , and  7  illustrate a nozzle cooling arrangement in which at least some of the FLADE airflow  80  is used as cooling air  251  which flowed through the hollow struts  208  into the substantially hollow centerbody  72 . The cooling air  251  is then flowed through cooling apertures  249  in the centerbody  72  downstream of the variable throat area A 8  to cool an outer surface of the centerbody. Some of the, FLADE airflow  80  may also be used as cooling air  251  for cooling the radially annular outer wall  220  of the engine nozzle  218  downstream of the variable throat area A 8  in the same manner. Cooling of the annular outer wall  220  and the hollow centerbody  72  is helpful when thrust augmenting afterburners  224 , such as those illustrated in  FIG. 7 , aft and downstream of the aft FLADE turbine  160  are ignited. The apertures may be angled to provide film cooling along the centerbody  72  and/or the hollow struts  208 . Holes, shaped and angled holes, and slots and angled slots are among the types of cooling apertures  249  that may be used. 
   The augmenter includes an exhaust casing  231  and liner  234  within which is defined a combustion zone  236 . The thrust augmenting afterburner  224  is mounted between the turbines and the exhaust nozzle for injecting additional fuel when desired during reheat operation for burning in the augmenter for producing additional thrust. In a bypass turbofan engine, an annular bypass duct extends from the fan to the augmenter for bypassing a portion of the fan air around the core engine to the augmenter. The bypass air is used in part for cooling the exhaust liner and also is mixed with the core gases prior to discharge through the exhaust nozzle. 
   Various types of flameholders are known and typically include radial and circumferential V-shaped gutters which provide local low velocity recirculation and stagnation regions therebehind, in otherwise high velocity core gas flow, for sustaining combustion during reheat operation. Since the core gases are the product of combustion in the core engine, they are initially hot when they leave the turbine, and are further heated when burned with the bypass air and additional fuel during reheat operation. 
   Illustrated in  FIGS. 3 and 4  is an embodiment of the engine  1  in which the fan inlet  11  to the fan section  115  is axially offset from the annular FLADE inlet  8  to the FLADE duct  3 . The exemplary axially offset FLADE inlet  8  is illustrated as being axially located substantially aftwardly of the fan section  115  and, more particularly, it is axially located aftwardly of the core engine  18 . An aircraft  124  illustrated in  FIG. 4  includes a single offset flush mounted engine air intake  127  connected to and in fluid communication with the annular fan inlet  11  by an engine fixed inlet duct  126 . Flush mounted dual FLADE air intakes  129  are connected to and in fluid communication with the annular FLADE inlets  8  by FLADE fixed inlet ducts  128 . The FLADE air intakes  129  are axially offset from the engine air intake  127 . This provides great flexibility in designing and constructing efficient engines, aircraft, and aircraft with engines completely mounted within the aircraft&#39;s fuselage  113  or body and the FLADE air intakes  129  and the engine air intake  127  are mounted flush with respect to the fuselage  113 . Also note that the FLADE duct  3  leads to separate FLADE exhaust nozzles  125  that are offset from the engine nozzle  218  which may be a variable or fixed throat area engine nozzle. Inlet duct passages  111  of the engine and FLADE fixed inlet ducts  126  and  128  may be two-dimensional terminating in transition sections  119  between the inlet duct passages  111  and the axisymmetric annular fan and FLADE inlets  11  and  8 . The FLADE airflow  80  may be also used as a heat sink for aircraft systems cooling purposes because it is a good source of pressurized cooling air. An example of this feature is illustrated as an aircraft waste heat cooling system  152  having an aircraft waste heat source  153  cooled by Flade exhaust airflow  154  in at least one FLADE air exhaust duct  155  leading from the FLADE duct  3  to one of the FLADE exhaust nozzles  125 . A heat exchanger  156  is positioned in the FLADE air exhaust duct  155  so it can be cooled by the pressurized Flade exhaust air  154 . Hot and cold fluid pipes  157  and  158  respectively (ducts may be used for air to air heat exchangers) fluidly connect the waste heat source  153  to the heat exchanger  156 . 
   Illustrated in  FIG. 5  is a variation of the embodiment of the aft FLADE engine  1  illustrated in  FIG. 3 . A “bolt on” aft FLADE module  260  incorporates a free aft FLADE turbine  160  and can be added to an existing engine  262  for various purposes including but not limited to testing and design verification. Another feature illustrated in  FIG. 5  is a FLADE power extraction apparatus  264 , illustrated as an electrical generator  266  disposed within the engine  1  and drivenly connected through a speed increasing gearbox  268  to the aft FLADE turbine  160 . The electrical generator  266  is illustrated as being located within the hollow engine nozzle centerbody  72  but may be placed elsewhere in the engine  1  as illustrated in  FIG. 11 . Another embodiment of the power extraction apparatus  264  is a power takeoff assembly  270 , as illustrated in  FIG. 9 , including a housing  274  disposed within the hollow engine nozzle centerbody  72 . A power takeoff shaft  276  is drivenly connected to the aft FLADE turbine  160  through a right angle gearbox  278  within the housing  274 . A power takeoff shaft is typically used to drive accessory machinery mounted external to the engine such as gearboxes, generators, oil and fuel pumps. The FLADE power extraction apparatus  264  allows more flexibility in the design of the engine  1  so that the power used by the aft FLADE fan blades  5  is a small percentage of the power extracted by the aft FLADE turbine  160  from the mixed flow  188  and, therefore, varying and controlling the amount of the FLADE airflow  80  will have a small effect on the efficiency of the aft FLADE turbine  160 . Also illustrated in  FIG. 5 , as well as in  FIGS. 9 and 10 , is a fixed throat area engine nozzle  216  axially aft of the mixer  49  and the fan bypass duct  40 . 
   Illustrated in  FIG. 10  is an engine  1  with the FLADE inlet  8  and the fan inlet  11  axially located together and not axially offset from each other as the embodiments illustrated in  FIGS. 3 and 5  and having the electrical generator  266  disposed within the hollow engine nozzle centerbody  72  and drivingly connected through the speed increasing gearbox  268  to the aft FLADE turbine  160 . Also note that the engines  1  illustrated in  FIGS. 5–7  and  9 – 10  are of the single bypass type having but a single bypass inlet  272  as compared to the engines  1  illustrated in  FIGS. 1–3  having both the first and second bypass inlets  42  and  46  to the fan bypass duct  40 . 
   Illustrated in  FIG. 11  is a portion of an engine  1  with more than one FLADE power extraction apparatus  264  disposed within the engine  1 . Forward and aft electrical generators  366  and  367  are disposed within the engine  1  forward and aft or downstream and upstream of the aft FLADE turbine  160 . The forward and aft electrical generators  366  and  367  are drivenly connected through forward and aft speed increasing gearboxes  368  and  369  to the aft FLADE turbine  160 . Also illustrated in  FIGS. 9–11  is a fixed throat area engine nozzle  218  having a fixed throat area A 8  downstream and axially aft of the aft FLADE turbine  160 . Power extraction may be accomplished in such a fixed throat area engine with the variable first FLADE vanes  6  scheduled closed. 
   Illustrated in  FIG. 6  is an engine  1  with a forward afterburner  226  axially disposed in the mixed flow  188  between the mixer  49  and the aft FLADE turbine  160 . The forward afterburner  226  includes forward fuel spraybars  230  and forward flameholders  232 . The forward afterburner  226  may be used to add additional energy to the mixed flow  188  upstream of the aft FLADE turbine  160  if more power is required for the aft FLADE turbine  160  to provide additional energy upon demand to the aft FLADE turbine  160  for the aft FLADE fan blades  5  and/or the power extraction apparatus  264  such as the electrical generator  266  or the power takeoff assembly  270 . 
   Schematically illustrated in cross-section in  FIG. 8  is an aircraft aft FLADE engine  1  having a fan section  115  with first and second counter-rotatable fans  130  and  132 . The variable first FLADE vanes  6  are used to control the amount of a FLADE airflow  80  allowed into the FLADE inlet  8  and the FLADE duct  3 . Opening of the FLADE duct  3  by opening the first FLADE vanes  6  at part power thrust setting of the FLADE engine  1  allows the engine to maintain an essentially constant inlet airflow over a relatively wide range of thrust at a given set of subsonic flight ambient conditions such as altitude and flight Mach No. and also avoid spillage drag and to do so over a range of flight conditions. This capability is particularly needed for subsonic part power engine operating conditions. 
   The FLADE inlet  8  and the fan inlet  11  in combination generally form the engine inlet  13 . Downstream and axially aft of the first and second counter-rotatable fans  130  and  132  is the core engine  18  having an annular core engine inlet  17  and a generally axially extending axis or centerline  12  generally extending forward  14  and aft  16 . A fan bypass duct  40  located downstream and axially aft of the first and second counter-rotatable fans  130  and  132  circumscribes the core engine  18 . The FLADE duct  3  circumscribes the first and second counter-rotatable fans  130  and  132  and the fan bypass duct  40 . 
   One important criterion of inlet performance is the ram recovery factor. A good inlet must have air-handling characteristics which are matched with the engine, as well as low drag and good flow stability. For a given set of operating flight conditions, the airflow requirements are fixed by the pumping characteristics of the FLADE engine  1 . During supersonic operation of the engine, if the area of the engine inlet  13  is too small to handle, the inlet airflow the inlet shock moves downstream of an inlet throat, particularly if it is a fixed inlet, and pressure recovery across the shock worsens and the exit corrected flow from the inlet increases to satisfy the engine demand. If the FLADE engine inlet area is too large, the engine inlet  13  will supply more air than the engine can use resulting in excess drag (spillage drag), because we must either by-pass the excess air around the engine or “spill” it back out of the inlet. Too much air or too little air is detrimental to aircraft system performance. The FLADE fan  2  and the FLADE duct  3  are designed and operated to help manage the inlet airflow delivered by the inlet to the fans. 
   The core engine  18  includes, in downstream serial axial flow relationship, a core driven fan  37  having a row of core driven fan blades  36 , a high pressure compressor  20 , a combustor  22 , and a high pressure turbine  23  having a row of high pressure turbine blades  24 . A high pressure shaft  26 , disposed coaxially about the centerline  12  of the engine  1 , fixedly interconnects the high pressure compressor  20  and the high pressure turbine blades  24 . The core engine  18  is effective for generating combustion gases. Pressurized air from the high pressure compressor  20  is mixed with fuel in the combustor  22  and ignited, thereby, generating combustion gases. Some work is extracted from these gases by the high pressure turbine blades  24  which drives the core driven fan  37  and the high pressure compressor  20 . The high pressure shaft  26  rotates the core driven fan  37  having a single row of circumferentially spaced apart core driven fan blades  36  having generally radially outwardly located blade tip sections  38  separated from generally radially inwardly located blade hub sections  39  by an annular fan shroud  108 . 
   The combustion gases are discharged from the core engine  18  into a low pressure turbine section  150  having counter-rotatable first and second low pressure turbines  19  and  21  with first and second rows of low pressure turbine blades  28  and  29 , respectively. The second low pressure turbine  21  is drivingly connected to the first counter-rotatable fan  130  by a first low pressure shaft  30 , the combination or assembly being designated a first low pressure spool  240 . The first low pressure turbine  19  is drivingly connected to the second counter-rotatable fan  132  by a second low pressure shaft  31 , the combination or assembly being designated a second low pressure spool  242 . The second counter-rotatable fan  132  has a single row of generally radially outwardly extending and circumferentially spaced-apart second fan blades  32 . The first counter-rotatable fan  130  has a single row of generally radially outwardly extending and circumferentially spaced-apart first fan blades  33 . The FLADE fan blades  5  are primarily used to flexibly match inlet airflow requirements. 
   The high pressure turbine  23  includes a row of high pressure turbine (HPT) nozzle stator vanes  110  which directs flow from the combustor  22  to the row of high pressure turbine blades  24 . Flow from the row of high pressure turbine blades  24  is then directed into counter-rotatable second and first low pressure turbines  21  and  19  and second and first rows of low pressure turbine blades  29  and  28 , respectively. A row of low pressure stator vanes  66  are disposed between the second and first rows of low pressure turbine blades  29  and  28 . 
   A row of fixed low pressure stator vanes  66  is disposed between the second and first rows of low pressure turbine blades  29  and  28 . Alternatively, a row of variable low pressure stator vanes may be incorporated between the second and first rows of low pressure turbine blades  29  and  28 . The first low pressure turbine  19  and its first row of low pressure turbine blades  28  are counter-rotatable with respect to the row of high pressure turbine blades  24 . The first low pressure turbine  19  and its first row of low pressure turbine blades  28  are counter-rotatable with respect to the second low pressure turbine  21  and its second row of low pressure turbine blades  29 . The aft FLADE turbine  160  is illustrated, in  FIG. 8  as a free turbine not connected to a spool or fan in the fan section  115 . Alternatively, the aft FLADE turbine  160  may be drivingly connected to the second low pressure shaft  31  of the second low pressure spool  242 . 
   The engines illustrated herein are single and double bypass types and it is thought that a turbojet type may be used in which there is no bypass duct or bypass flow and the aft FLADE turbine would be placed downstream of any turbine section used to drive the fan and/or compressor. Turbojet type engines may also use augmenters and variable area two-dimensional nozzles. 
   While there have been described herein what are considered to be preferred and exemplary embodiments of the present invention, other modifications of the invention shall be apparent to those skilled in the art from the teachings herein and, it is therefore, desired to be secured in the appended claims all such modifications as fall within the true spirit and scope of the invention. Accordingly, what is desired to be secured by Letters Patent of the United States is the invention as defined and differentiated in the following claims.