Abstract:
A case for a gas turbine engine includes a core body. The core body defines a longitudinally extending core flow path, a laterally extending bleed air duct coupling the core flow path in fluid communication with the external environment, and a structure-supporting member spanning the bleed air duct. A heating element is connected to the core body and is in thermal communication with the structure-supporting member.

Description:
CROSS REFERENCE TO RELATED APPLICATIONS 
       [0001]    This application claims the benefit of U.S. Provisional Patent Application Ser. 62/092,046 filed on Dec. 15, 2014 the entire contents of which are incorporated herein by reference thereto. 
     
    
     BACKGROUND OF THE DISCLOSURE 
       [0002]    1. Field of the Disclosure 
         [0003]    The present disclosure relates to gas turbine engines, and more particularly to heated cases housing gas turbine engine rotating components. 
         [0004]    2. Description of Related Art 
         [0005]    Gas turbine engines commonly include a compressor section with two or more compressor stages ordinarily sealed from the external environment. Under certain circumstances, it can become necessary to bleed compressed air from the compressor section to the external environment, typically through bleed air ducts defined through the case housing the compressor section. This can be necessary to adjust airflow and pressure ratio of fluid traversing the compressor section stages. Bleeding compressor airflow can also allow foreign material ingested by the compressor section, such as rain, ice, or hail, to be extracted from air traversing the compressor stages. 
         [0006]    Some engine cases include ducts arranged about an annulus of the case. Valves coupled to the ducts selectively place the core interior in fluid communication with the external environment for bleeding airflow from the compressor section. Under certain operational conditions, portions of the case bounding the bleed air ducts can collect foreign material extracted from the core flow path, potentially blocking such ducts or causing the foreign material to be passed back to the compressor. 
         [0007]    Such conventional methods and systems have generally been considered satisfactory for their intended purpose. However, there is still a need in the art for improved gas turbine engine cases. The present disclosure provides a solution for this need. 
       SUMMARY OF THE DISCLOSURE 
       [0008]    A case for a gas turbine engine includes a core body. The core body defines a longitudinally extending core flow path, a laterally extending bleed air duct coupling the core flow path in fluid communication with the external environment, and a structural member spanning the bleed air duct. A heating element connects to the core body and is in thermal communication with the structural member. 
         [0009]    In certain embodiments, the heating element can be disposed within a bore defined within the core body. The bore can be defined within in a forward segment, an aft segment, or the structural member of the core body. The bore can have an orientation with a longitudinal, lateral or circumferential component relative to engine rotation axis within the core body. The heating element can include a resistive heating element, such as a cartridge type heating element, and can seat within the bore such that it is in thermal communication with a core flow path-facing surface of the structural member. 
         [0010]    It also contemplated that, in accordance with certain embodiments, the heating element can overlay at least a portion of the core body. The heating element can overlay an exterior surface of the forward segment, the aft segment, of an exterior surface of the bleed air duct extending radially outward from the core body. The heating element can overlay an interior surface of forward segment, the aft segment, or structural member of the core body. The heating element can also overlay an inlet portion of the bleed air duct. It is further contemplated that the heating element can be thermal communication with a core flow path-facing surface of the structural member, and can include a heater mat type heating element. 
         [0011]    It is further contemplated that, in certain embodiments, the bleed air duct can be an annulus dividing the core body into a forward segment and an aft segment. The structural member can span the annulus and couple the forward segment to the aft segment. The structural member can also divide the annulus into a plurality of circumferentially adjacent bleed air ducts. It is also contemplated that the structural member can also have a structural member surface with an aerodynamic surface facing the core flow path. 
         [0012]    A system for heating a case for a gas turbine engine includes a case with a core body as described above, a processor operatively associated with the heating element, and a memory communicative with the processor. The memory has instructions recorded on the memory that, when read by the processor, cause the processor to determine the flight condition of an aircraft, compare the flight condition to a programmed condition, and change the amount of electrical power applied to the heating element, comparing the flight condition to the programmed condition operation to determine that the programmed flight condition is met. 
         [0013]    A method of heating a gas turbine engine case includes applying electrical power to a heating element disposed on or within a body of an engine case. The method additionally includes heating, using the heating element, an interior surface of the body bounding a flow path disposed within the engine. 
         [0014]    In embodiments, the method can include continuously heating the surface. In certain embodiments, the method can include selectively heating the surface. Selectively heating the surface can include determining a flight condition of an aircraft and comparing the flight condition to a programmed condition. When the comparison indicates that the programmed condition has been met, an amount electrical power applied to a heating element can be changed. 
         [0015]    It is also contemplated that, in accordance with certain embodiments, the flight condition can be when hail or ice ingestion can be expected. The method can also include changing the amount of power can include increasing power applied when the programmed flight condition is met. The method can further include decreasing power applied when the programmed flight condition is not met. 
         [0016]    These and other features of the systems and methods of the subject disclosure will become more readily apparent to those skilled in the art from the following detailed description of the preferred embodiments taken in conjunction with the drawings. 
     
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         [0017]    So that those skilled in the art to which the subject disclosure appertains will readily understand how to make and use the devices and methods of the subject disclosure without undue experimentation, preferred embodiments thereof will be described in detail herein below with reference to certain figures, wherein: 
           [0018]      FIG. 1  is a schematic partial cross-sectional view of an exemplary embodiment of a gas turbine engine constructed in accordance with the present disclosure, showing a core case; 
           [0019]      FIG. 2  is a cross-sectional view of a portion the gas turbine engine of  FIG. 1 , showing a bleed air duct extending from the core case interior to the external environment; 
           [0020]      FIG. 3  is a cross-sectional view of a portion of the gas turbine engine of  FIG. 1 , showing cartridge type heating elements disposed within the core case bout the bleed air duct; and 
           [0021]      FIG. 4  is a cross-sectional view of a portion of the gas turbine engine of  FIG. 1 , showing heater mat type heating elements disposed over surfaces of the core case about the bleed air duct. 
       
    
    
     DETAILED DESCRIPTION OF THE EMBODIMENTS 
       [0022]    Reference will now be made to the drawings wherein like reference numerals identify similar structural features or aspects of the subject disclosure. For purposes of explanation and illustration, and not limitation, a partial view of an exemplary embodiment of a core case in accordance with the disclosure is shown in  FIG. 1  and is designated generally by reference character  100 . Other embodiments of core cases in accordance with the disclosure, or aspects thereof, are provided in  FIGS. 2-4 , as will be described. The systems and methods described herein can be used for aero and industrial gas turbine engines, such as aircraft main engines or auxiliary power units, and in power plants for electricity generation. 
         [0023]      FIG. 1  schematically illustrates a gas turbine engine  20 . 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 augmenter section (not shown) among other systems or features. Although gas turbine engine  20  is depicted as a geared 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 geared turbofans as the teachings may be applied to other types of turbine engines including three-spool turbofan engines or geared turbofans, or turboshaft engines. 
         [0024]    Fan section  22  drives air along a bypass flow path B while compressor section  24  drives air along a core flow path C for compression and communication into the combustor section  26  and expansion through the turbine section  28 . Gas turbine engine  20  generally includes a low speed spool  30  and a high speed spool  32  mounted for rotation about an engine rotation axis R relative to an engine core case  100  via several bearing systems  38 . It should be understood that various bearing systems  38  at various locations may alternatively or additionally be provided. 
         [0025]    Low speed spool  30  generally includes an inner shaft  40  that interconnects a fan  42 , a low pressure compressor  44  and a low pressure turbine  46 . Inner shaft  40  is connected to fan  42  through a geared architecture  48  to drive the fan  42  at a lower speed than low speed spool  30 . Geared architecture  48  connects the low pressure compressor  44  to fan  42 , but allows for rotation of low pressure compressor  44  at a different speed and/or direction than fan  42 . 
         [0026]    High speed spool  32  includes an outer shaft  50  that interconnects a high pressure compressor  52  and a high pressure turbine  54 . A combustor  56  is arranged between high pressure compressor  52  and the high pressure turbine  54 . A mid-turbine frame  58  disposed with engine core case  100  is arranged generally between the high pressure turbine  54  and the low pressure turbine  46 . The mid-turbine frame  58  further supports bearing systems  38  in turbine section  28 . 
         [0027]    Inner shaft  40  and outer shaft  50  are concentric and rotate via bearing systems  38  about the engine rotation axis R that is collinear with their respective longitudinal axes. Core airflow C is compressed by low pressure compressor  44  then 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 . High pressure turbine  54  and low pressure turbine  46  rotationally drive the respective low speed spool  30  and high speed spool  32  in response to the expansion. 
         [0028]    With reference to  FIG. 2 , a portion of low pressure compressor  44  is shown. Low pressure compressor  44  includes a rotor stage  60  and a stator stage  62  housed within engine core case  100 . Rotor stage  60  is forward of stator stage  62  and upstream relative to core airflow C. Engine core case  100  defines a bleed air duct  64  between rotor stage  60  and stator stage  62 . Bleed air duct  64  extends radially outward relative core flow path C and engine rotation axis R. A valve assembly  66  (shown schematically) is connected to bleed air duct  64  on a radially outer end of bleed air duct  64  that is configured and adapted to bleed air from between rotor stage  60  and stator stage  62  under predetermined conditions, such as matching airflow through low pressure compressor  44  and high pressure compressor  52  or extracting foreign material from core flow path C. In this respect valve assembly  66  includes a movable member (not shown for clarity reasons) with a first position, wherein substantially all air traversing low pressure compressor  44  along core flow path C is communicated to high pressure compressor  52  (shown in  FIG. 1 ), and a second position wherein at least a portion of air traversing low pressure compressor  44  is communicated as a bleed airflow D to the environment external to engine core case  100  through bleed air duct  64 . 
         [0029]    With reference to  FIG. 3 , engine core case  100  is shown. Engine core case  100  has a forward segment  102  and an aft segment  104  that define there between an annulus  110 . Annulus  110  leads to a bleed air duct  64  and is spanned by a structure-supporting member  106 . Structure-supporting member  106  couples forward segment  102  to aft segment  104 . In embodiments, structure-supporting member  106  bounds a pair of circumferentially adjacent bleed air ducts. In certain embodiments, core case structure-supporting member  106  includes an aerodynamic surface bounding an interior of the core body. 
         [0030]    An aft-facing edge of forward core case segment  102 , core flow path-facing surface  108 , and forward-facing edge of aft core case segment  104  bound an inlet of bleed air duct  64 . As illustrated, bleed air duct  64  is located at an axial engine station disposed between low pressure compressor  44  and high pressure compressor  52  (shown in  FIG. 1 ), i.e. a 2.5 bleed duct. 
         [0031]    During operation in hail events valve assembly  66  can be opened to extract hail ingested by gas turbine engine  20 . In this regard opening valve assembly  66  generates a bleed airflow D that flows through bleed air duct  64 . Bleed airflow D extracts foreign material traversing compressor section  24  along core flow path C through bleed air duct  64  and into the environment external to gas turbine engine  20 . Hail impinging a core flow path-facing surface  108  of structural member  106  can lower the temperature of the surface. The temperature drop can be sufficient such that hail and/or ice accumulate on core flow path-facing surface  108  instead of exiting the case through bleed air duct  64 . Engine operating conditions can lower the temperature of the surface sufficient such that hail and/or ice can accumulate on the core flow path-facing surface  108 . Under certain circumstances, accumulated ice and/or hail can also be returned to core flow path C. 
         [0032]    Engine core case  100  includes one or more bores having one or more heating elements seated therein for heating core flow path-facing surface  108 , thereby making it more difficult for ice and/or hail to accumulate on core flow path-facing surface  108 . In this respect, core case structure-supporting member  106  defines a structural member bore  120  seating a heating element  122 . Structural member bore  120  can have an orientation with an axial component relative to engine rotation axis R, for example being angled in relation thereto, or can be substantially parallel in relation to engine rotation axis R. This positions heating element  122  axially and substantially in parallel with core flow path-facing surface  108 . It is to be understood and appreciated that bore  120  (and the cartridge type heating element seated therein) can have an orientation with a longitudinal, lateral, radial and/or a circumferential component relative to engine rotation axis R as suitable for an intended application for heating core flow path-facing surface  108 . 
         [0033]    Alternatively or additionally, core aft segment  104  also defines an aft segment bore  130  seating a heating element  132 . Aft segment bore  130  is oriented radially relative to engine rotation axis R (shown in  FIG. 1 ). The radially orientation positions heating element  132  substantially orthogonal relative to core flow path-facing surface  108 . As above, bore  130  (and the cartridge-type heating element seated therein) can have an orientation with a longitudinal, lateral, radial and/or a circumferential component relative to engine rotation axis R as suitable for an intended application for heating core flow path-facing surface  108 . It also to be appreciated and understood that forward segment  102  can also define a bore seating a cartridge-type heating element. Heating element  122  and/or heating element  132  can include a cartridge-type heating elements. In yet another alternative embodiment, the heating element or elements can, alternatively or additionally, be arranged such that it is located within another location of the bleed air duct  64  for example, proximate to an interface of a bleed air duct located within the fan bypass duct that exhausts to the atmosphere or to another region of the engine. Examples of suitable cartridge heater include OMEGALUX® CIR cartridge heaters, available from Omega Engineering Inc. of Stamford, Conn. 
         [0034]    With reference to  FIG. 4 , an engine core case  200  is shown. Engine core case  200  is similar to engine core case  100  and additionally includes at least one surface heating element. At least one heating element can include a conformal heating element. The surface heating element can be arranged on an external or interior surface of core case  200 . For example, a heating element  210  can be arranged on an external surface  212  of aft segment  204 . Alternatively (or additionally), a heating element  220  can be arranged on interior surface  222  of aft segment  204 . A heating element  232  can also be arranged on core path-facing surface  208  of structure-supporting member  206 . As will be appreciated, surface heating elements can be arranged on interior and/or exterior surfaces of forward segment  202  as well as on an interior surface of bleed air duct  64 , potentially heating directly surfaces susceptible to ice accumulation. In yet another alternative embodiment, the heating element or elements can, alternatively or additionally, be arranged such that it is located within another location of the bleed air duct  64  for example, proximate to an interface of a bleed air duct located within the fan bypass duct that exhausts to the atmosphere or to another region of the engine. Examples of suitable heating elements include OMEGALUX® silicone rubber flexible heaters, also available from Omega Engineering Inc. of Stamford, Conn. 
         [0035]    With reference to  FIG. 1 , system  300  for heating a gas turbine engine core case is shown. System  300  includes a processor  302  and a memory  304 . Processor  302  is operatively associated with heating element  122  and communicative with a memory  304 . Memory  304  has instructions recorded thereon that, when read by processor  302 , cause processor  302  to undertake certain actions. In particular, the instructions cause processor  302  to determine a flight condition of an aircraft, compare the flight condition of the aircraft to a programed condition, and change the amount of power supplied to a heating element connected to core engine case  100  or engine core case  200  when the comparison indicates that the programmed flight condition is met. 
         [0036]    For example, if the comparison indicates that an aircraft-mounted gas turbine engine  20  is beginning a descent from altitude to landing, processor  302  can increase power provided to the heating element to reduce the risk of hail or ice accumulations on core flow path-facing surface  108  (shown in  FIG. 3 ). Alternatively, if the comparison indicated an aircraft-mounted gas turbine engine  20  is operating in icing or hailing conditions, processor  302  can increase power provided to the heating element to reduce the risk of hail or ice accumulation on core flow path-facing surface  108  (shown in  FIG. 3 ). This can reduce or substantially eliminate accumulation of ice particles on core case structure-supporting members, allowing bleed air duct  64  to more efficiently extract hail from the core flow of path gas turbine engine  20 . Heating core flow path-facing surface  106  can allow for operation of gas turbine engine  20  at power settings (i.e. rotation speeds) more favorable to engine efficiency rather than speeds favorable for hail extraction. It is further contemplated that the instructions can cause processor  302  to provide reduced, or substantially no power, to the element(s) when the comparison indicates that it is unlikely that gas turbine engine  20  will encounter hail. 
         [0037]    The methods and systems of the present disclosure, as described above and shown in the drawings, provide for gas turbine engines with superior properties including improved efficiency during operation in environments where hail can be encountered. While the apparatus and methods of the subject disclosure have been shown and described with reference to preferred embodiments, those skilled in the art will readily appreciate that changes and/or modifications may be made thereto without departing from the scope of the subject disclosure.