Patent Publication Number: US-8967543-B2

Title: Aircraft ice protection system and aircraft provided with the same

Description:
TECHNICAL FIELD 
     The present invention relates to an aircraft ice protection system for preventing ice accretion on the outside of a wing leading edge of an aircraft or removing the accreted ice during flight, and also relates to an aircraft provided with the ice protection system. 
     This application claims the right of priority to Japanese Patent Application No. 2010-192618 filed in Japan on Aug. 30, 2010, the content of which is hereby incorporated herein. 
     BACKGROUND ART 
     An aircraft is provided with ice protection equipment for preventing ice accretion on a wing leading edge of the aircraft or removing the accreted ice during flight. Then, the ice protection equipment is provided with a hot air chamber formed inside the wing leading edge of a main wing, and bleed air (high-temperature air) extracted from a main engine is supplied to the hot air chamber to heat the wing leading edge of the main wing from inside (refer to Patent Document 1 and Patent Document 2, for example). 
     PRIOR ART DOCUMENTS 
     Patent Documents 
     
         
         [Patent Document 1] Japanese Published Unexamined Utility Model Application No. H1-149894 
         [Patent Document 2] Japanese Published Unexamined Patent Application No. H6-206593 
       
    
     SUMMARY OF THE INVENTION 
     Problems to be Solved by the Invention 
     In general, the above-described conventional ice protection equipment is operated only under conditions where ice accretion easily takes place on an airframe, for example, in clouds or the like. However, since a wing leading edge is heated with bleed air extracted from a main engine of an aircraft, problems are posed such as a decrease in the original impelling force of a main engine and poor operation cost associated with a decrease in fuel consumption efficiency. That is, these problems partially contribute to reduced performance of an aircraft. 
     In view of the above-described situation, the present invention provides an aircraft ice protection system which is capable of effectively carrying out ice protection and also capable of improving the performance of an aircraft such as suppression of a decrease in impelling force of a main engine and an improvement in fuel consumption efficiency, and also provides an aircraft provided with the ice protection system. 
     Means for Solving the Problems 
     In order to attain the above-described object, the invention provides the following means. 
     The aircraft ice protection system of the present invention is an ice protection system for preventing ice accretion on a wing of an aircraft or removing the accreted ice. The ice protection system is provided with a hot air chamber which is formed inside the wing of the aircraft, a bleed air supply line which supplies bleed air extracted from a main engine of the aircraft to the hot air chamber, a heated air supply line which supplies air introduced from an air intake of the aircraft to the hot air chamber via a heat source of the aircraft, and switching device which selectively supplies the bleed air and the air heated by the heat source to the hot air chamber. The wing is heated by the air supplied to the hot air chamber, and ice protection of the wing is carried out. 
     In this invention, as with a conventional invention, the bleed air is supplied from the main engine to the hot air chamber, thereby protecting the wing (wing leading edge) from ice, the air introduced from the air intake installed on the airframe is also fed into a heat source of the airframe of the aircraft by ram pressure, and hot air (heated air) is produced by utilizing the heat source. Then, ice protection is carried out by supplying the heated air to the hot air chamber. 
     Further, in the aircraft ice protection system of the present invention, it is preferred that the heat source be at least one of an oil cooler, an oil tank, and a main engine of the aircraft, the oil cooler and the oil tank being mounted on a hydraulic circuit provided on the aircraft. 
     In this invention, as a heat source for heating the air introduced from the air intake, the oil cooler or the oil tank on the hydraulic circuit or the main engine of the aircraft is adopted, thereby reliably heating the air at such a temperature that can be utilized for ice protection. 
     In particular, conventionally, air introduced from the air intake is used as a cooling medium and heated air after heat exchange is discharged outside. However, in this invention, the heated air is effectively used for ice protection. 
     Further, when heat exchange is performed with hydraulic oil pooled in the oil tank to heat the air, the present invention performs not only cooling the hydraulic oil, which is at a high temperature, by an oil cooler but also causing temperature decrease (making low in temperature) of the hydraulic oil inside the oil tank by exchanging heat with the air. Thereby, the hydraulic oil can be delayed in oxidation and deterioration, and replacement interval of the hydraulic oil is prolonged. 
     Still further, in the aircraft ice protection system of the present invention, it is preferred that the oil cooler be provided with a double-pipe structured heat transfer pipe which is composed of an inner pipe through which hydraulic oil flows and an outer pipe through which the air introduced from the air intake flows between the outer pipe and the inner pipe, and the air be heated by exchanging heat with the hydraulic oil flowing through the inner pipe. 
     In this invention, the oil cooler is configured in such a manner that the air introduced from the air intake is allowed to flow between the outer pipe and the inner pipe of the double-pipe structured heat transfer pipe. Thereby, heat exchange is performed with the hydraulic oil flowing through the inner pipe to reliably and efficiently heat the air, and the heated air is effectively used for ice protection. 
     Further, in the aircraft ice protection system of the present invention, it is further preferred that projected portions be provided on an outer face of the inner pipe. 
     In this invention, the projected portions are provided on the outer face of the inner pipe to increase a heat transfer area. Further, when the air flows between the outer pipe and the inner pipe, the flow of the air turns into a turbulent state at the projected portions, thereby facilitating heat exchange between the hydraulic oil and the air. As a result, the heat exchange is performed at a higher efficiency, and the piping length inside the oil cooler (the heat transfer pipe length) is decreased. Further, the piping inside the oil cooler can be decreased in length as described above, which contributes to a reduction in weight of the airframe of the aircraft. 
     Still further, in the aircraft ice protection system of the present invention, it is preferred that the oil tank be provided with an outer shell which forms a flowing space between the outer shell and an outer face of a tank main body which pools hydraulic oil, and the air flowing through the flowing space be heated by exchanging heat with the hydraulic oil inside the tank main body. 
     In this invention, the oil tank is formed so as to have the outer shell which forms the flowing space between the outer shell and the outer face of the tank main body, thereby allowing the air to flow through the flowing space and reliably heating the air by using the oil tank as a heat source. 
     Further, in the aircraft ice protection system of the present invention, it is preferred that projected portions be provided on an outer face of the tank main body. 
     In this invention, since the projected portions are provided on the outer face of the tank main body, a heat transfer area is increased. Further, when the air flows through the flowing space between the tank main body and the outer shell, the flow the air turns into a turbulent state at the projected portions, thereby facilitating heat exchange with the hydraulic oil. Thus, the heat exchange is performed at a higher efficiency to reliably and efficiently heat the air, and the heated air is also effectively used for ice protection. 
     Still further, in the aircraft ice protection system of the present invention, it is preferred that the main engine be provided with an outer shell which forms a flowing space between the outer shell and an outer face of a casing, and the air be heated by allowing the air flowing through the flowing space. 
     In this invention, the main engine is configured so as to have an outer shell which forms a flowing space between the outer shell and, for example, an outer face of a casing of a burner, which is at a relatively high temperature. Air is allowed to flow through the flowing space, by which the air is heated, with the main engine used as a heat source. 
     Further, in the aircraft ice protection system of the present invention, it is preferred that projected portions be provided on the outer face of the casing. 
     In this invention, since the projected portions such as the fins are provided on the outer face of the casing, a heat transfer area is increased. Further, when the air flows through the flowing space between the casing and the outer shell, the flow of the air turns into a turbulent state at the projected portions, thereby facilitating heat exchange with the main engine. As a result, the heat exchange is performed at a higher efficiency to reliably and efficiently heat the air, and the heated air is also effectively used for ice protection. 
     Further, in the aircraft ice protection system of the present invention, it is preferred that the projected portions on the outer face of the casing be formed in such a manner that the air flows while swirling along the outer face of the casing. 
     In this invention, the air is allowed to flow while swirling along the outer face of the casing by the projected portions on the outer face of the casing. Thereby, contact time of the air on the casing is increased to raise the efficiency of heat exchange, thus making it possible to reliably and efficiently heat the air and also effectively use the heated air for ice protection. 
     In the aircraft ice protection system of the present invention, the switching device may include a first on-off valve installed on the bleed air supply line, a second on-off valve installed on the heated air supply line, and a controller for controlling operation of the first and the second on-off valves. 
     The aircraft of the present invention is provided with any one of the above-described aircraft ice protection systems. In this invention, working effects of the above-described aircraft ice protection systems are obtained. 
     Effects of the Invention 
     According to the aircraft ice protection system of the present invention and the aircraft provided with the ice protection system, in addition to a same configuration as the conventional which supplies bleed air from the main engine to the hot air chamber, air introduced from the air intake installed on the airframe of the aircraft is fed into a heat source of the airframe of the aircraft by ram pressure, by which hot air (heated air) produced by utilizing the heat source is supplied to the hot air chamber to carry out ice protection. 
     Thereby, the bleed air and the air heated by the heat source are selectively supplied to the hot air chamber to carry out ice protection. Therefore, when the air heated by the heat source is able to sufficiently carry out ice protection, it is no longer necessary to use the bleed air. Further, when only the air heated by the heat source is insufficient in carrying out ice protection, the bleed air is used to reliably carry out ice protection. Still further, even when the bleed air is used, the heated air is able to keep a wing heat in advance (able to heat the wing in advance), realizing a state in which ice accretion is less likely to take place, thus making it possible to suppress an extract amount of the bleed air to a minimum level. 
     As described above, the bleed air extracted from the main engine becomes unnecessary or an extract amount of the bleed air is suppressed to a minimum level during flight. Therefore, as compared with an aircraft having conventional ice protection equipment, it is possible to improve the performance of the aircraft such as suppression of a decrease in impelling force of the main engine and a reduction in operation cost associated with improvement in fuel consumption efficiency. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         FIG. 1  is a drawing which shows an aircraft according to an embodiment of the present invention. 
         FIG. 2  is a drawing which shows an aircraft ice protection system of a First Embodiment in the present invention. 
         FIG. 3  is a drawing which shows a main engine and also a configuration for supplying bleed air from the main engine to a hot air chamber of a wing. 
         FIG. 4  is a cross-sectional view of the hot air chamber of the wing taken along X 1 -X 1  in  FIG. 3 . 
         FIG. 5  is a drawing which shows a heat transfer pipe mounted on an oil cooler of the aircraft ice protection system of the First Embodiment in the present invention. 
         FIG. 6  is a drawing which shows an aircraft ice protection system of a Second Embodiment in the present invention. 
         FIG. 7  is a drawing which shows a part of an oil tank (tank main body, outer shell and flowing space) of the aircraft ice protection system of the Second Embodiment in the present invention. 
         FIG. 8  is a drawing which shows an aircraft ice protection system of a Third Embodiment in the present invention. 
         FIG. 9(   a ) is a drawing which shows a flowing space (casing, outer shell and projected portions) formed on a main engine of an aircraft ice protection system of the Third Embodiment in the present invention, and  FIG. 9(   b ) is a sectional view taken along an arrow A in  FIG. 9(   a ). 
     
    
    
     MODE FOR CARRYING OUT THE INVENTION 
     First Embodiment 
     Hereinafter, a description will be given of an aircraft ice protection system of a First Embodiment in the present invention and an aircraft provided with the ice protection system with reference to  FIG. 1  to  FIG. 5 . 
     First, as shown in  FIG. 1  and  FIG. 2 , an aircraft  1  of the present embodiment is provided with a main engine  3  installed on a main wing  2 , a hydraulic pump  25  which uses the main engine  3  as a driving source, a hydraulic circuit  5  actuated by the hydraulic pump  25 , and a manifold  5 ′ for controlling an actuator  4  which configures a part of the hydraulic circuit  5  and is installed on the main wing  2 . 
     As shown in  FIG. 3 , a gas turbine  6 , which is the main engine  3 , is provided with a fan casing  7  and a core engine casing  8 . The gas turbine  6  houses a fan  9  inside the fan casing  7 , and also the gas turbine  6  houses a compressor  10 , a burner  11  and a turbine  12  inside the core engine casing  8 . 
     The fan  9  is formed by attaching a plurality of fan blades  16  on an outer circumference of a rotating shaft  15 . The compressor  10  is provided with a low-pressure compressor  17  and a high-pressure compressor  18 . The turbine  12  is provided with a high-pressure turbine  19  and a low-pressure turbine  20  and disposed on the downstream side of the compressor  10 . Then, the rotating shaft  15  of the fan  9  is connected to the low-pressure compressor  17 , and the low-pressure compressor  17  is connected to the low-pressure turbine  20  by a first rotor shaft  21 . Further, the high-pressure compressor  18  is connected to the high-pressure turbine  19  by a cylindrical second rotor shaft  22  positioned on an outer circumference side of the first rotor shaft  21 . 
     Then, air introduced from the air intake port is compressed by passing through a plurality of compressor vanes and compressor blades (not illustrated) installed on the low-pressure compressor  17  and the high-pressure compressor  18  of the compressor  10 , thereby turned into compressed air which is high temperature and high pressure. Further, a predetermined fuel is supplied to the compressed air in the burner  11  and the fuel is burnt to produce a combustion gas which is high temperature and high pressure. The produced combustion gas passes through the plurality of turbine vanes and turbine blades (not illustrated) installed on the high-pressure turbine  19  and the low-pressure turbine  20  which configure the turbine  12 , thereby rotating and driving the turbine  12 . At this time, power of the low-pressure turbine  20  is transmitted to the fan  9  by the first rotor shaft  21 , and the fan  9  sends air, thereby producing an impelling force. 
     As shown in  FIG. 1  and  FIG. 2 , the hydraulic circuit  5  of the present embodiment supplies and discharges hydraulic oil to the actuator  4  for a flight control system such as a flap, an aileron, a rudder and landing equipment to drive the actuator  4 . The hydraulic circuit  5  is provided with a hydraulic pump  25 , an oil cooler (heat exchanger)  26  and an oil tank  27 . Further, for example, as shown in  FIG. 1 , the hydraulic pump  25  is driven by the main engine and is installed inside a main engine nacelle. In addition, the oil cooler  26  and the oil tank  27  are installed at a lower part of a body (airframe)  28 . 
     Then, in the hydraulic circuit  5 , as shown in  FIG. 1  and  FIG. 2 , hydraulic oil is supplied from the oil tank  27  to the actuator  4  by driving the hydraulic pump  25  and the actuator  4  is actuated. At the same time, the hydraulic oil (return oil) from the actuator  4  is fed back to the oil tank  27  through the oil cooler  26 . At this time, the oil cooler  26  is formed in such a manner that the heat-transfer pipe (hydraulic line)  30  meanders. Thus, low-temperature air S 1  introduced from the air intake  31  installed on the body  28  is used as a cooling medium to carry out heat exchange between the air S 1  and return oil, which is at a high-temperature of 70° C. to 80° C., for example, thereby cooling the return oil. Then, the return oil is fed back to the oil tank  27 . 
     On the other hand, in the aircraft ice protection system A of the present embodiment, as shown in  FIG. 2  and  FIG. 3 , bleed air (high-temperature air) S 2  extracted from the main engine  3  and air S 1 ′ introduced from the air intake  31  and heated by the oil cooler  26  (a heat source of an airframe of an aircraft) are selectively supplied to the hot air chamber  32  formed inside a wing leading edge  2   a  of the main wing  2  to carry out ice protection. 
     Here, in the present embodiment, as shown in  FIG. 3  and  FIG. 4 , in the wing leading edge  2   a  of the main wing  2 , a tip thereof is formed in a curved shape by jointing, for example, upper and lower curved plates, and a guide plate  33  is disposed inside the wing leading edge  2   a , with a predetermined clearance kept. As with the wing leading edge  2   a , the tip of the guide plate  33  is also formed in a curved shape by jointing, for example, upper and lower curved plates. In addition, the wing leading edge  2   a  is a part which includes the vicinity of the leading edge of the wing  2 . 
     Further, the wing leading edge  2   a  extends along the longitudinal direction of the main wing  2  (in the perpendicular direction on the plane of the page of  FIG. 4 ). The guide plate  33  has a predetermined length along the width direction of the aircraft  1  and the plurality of guide plates  33  are installed together in this direction. Still further, a partition wall  34  is provided inside the wing leading edge  2   a  along the longitudinal direction of the main wing  2  and also along the front-back direction of the main wing  2  (in the lateral direction in  FIG. 3 ). The partition wall  34  is formed in the longitudinal direction of the main wing  2 , with a predetermined clearance kept. In addition, each of the guide plates  33  is installed by keeping both ends thereof in contact with or fixed on end faces of the partition wall  34 . 
     The hot air chamber  32  is formed by being surrounded with the wing leading edge  2   a , the partition wall  35  at the rear thereof, and partition walls  34  on both sides. Further, a space between the wing leading edge  2   a  and the guide plate  33  is given as a hot air path  36 , and the hot air path  36  runs along the inside of the wing leading edge  2   a  and extends from the tip of the wing leading edge  2   a  to the rear thereof and is opened to the hot air chamber  32 . The hot air chamber  32  is provided with a discharge port which discharges air inside the hot air chamber  32  to the outside, whenever necessary. 
     Further, as shown in  FIG. 4 , inside the hot air chamber  32 , two duct pipes  40 ,  41  are disposed at the tip side of the wing leading edge  2   a  so as to be adjacent to the guide plate  33 . Each of the duct pipes  40 ,  41  is formed with both ends closed and extends so as to penetrate through the partition wall  34  along the longitudinal direction of the main wing  2 . Still further, the guide plate  33  is provided with an opening  33   a  at a position corresponding to the tip of the wing leading edge  2   a . Respectively on the duct pipes  40 ,  41 , injection holes  40   a ,  41   a  are formed which are opened to the front of the wing leading edge  2   a  at a position facing to the opening  33   a  formed on the guide plate  33 . Then, the duct pipes  40 ,  41  are configured so that the injection holes  40   a ,  41   a  are connected respectively via connecting pipes  42 ,  43  to the opening  33   a  of the guide plate  33 . At this time, the openings  33   a , the injection holes  40   a ,  41   a , and the connecting pipes  42 ,  43  are provided in a plural number on a single guide plate  33 , with a predetermined clearance kept. In addition, these two duct pipes  40 ,  41  are installed by connecting the injection holes  40   a ,  41   a  to the opening  33   a  of the guide plate  33  respectively via the connecting pipes  42 ,  43 . That is, they are installed individually by being communicatively connected to the hot air path  36 . 
     Then, as shown in  FIG. 2  to  FIG. 4 , in the aircraft ice protection system A of the present embodiment, one duct pipe  40  is connected to the compressor  10  of the main engine  3  by piping (bleed air supply line)  45 , and bleed air S 2  of the compressor  10  of the main engine  3  is supplied from the bleed air supply line  45  to the hot air path  36  (hot air chamber  32 ). 
     Further, the other duct pipe  41  is connected to the oil cooler  26  mounted on the hydraulic circuit  5  by piping (heated air supply line)  46 . At this time, in the present embodiment, as shown in  FIG. 2  and  FIG. 5 , the meandering heat transfer pipe  30  of the oil cooler  26  is formed in a double pipe structure by an inner pipe  47  in which hydraulic oil discharged from the actuator  4  flows and an outer pipe  48  in which air S 1  introduced from the air intake  31  is allowed to flow between the outer pipe  48  and the inner pipe  47 . Further, a plurality of rings are attached to the inner pipe  47  in an integrated manner, and these rings form projected portions  49  projecting toward an inner face of the outer pipe  48  from an outer face of the inner pipe  47 . 
     Further, the other duct pipe  41  is connected to a flowing space  50  in which the air S 1  between the inner pipe  47  and the outer pipe  48  of the heat transfer pipe  30  of the oil cooler  26  flows via the heated air supply line  46 . 
     Still further, as shown in  FIG. 2 , the bleed air supply line  45  and the heated air supply line  46  are provided respectively with on-off valves  51 ,  52 . Then, in the aircraft ice protection system A of the present embodiment, these on-off valves  51 ,  52  are controlled for opening and closing by a valve controller  54  based on results measured by a thermometer  53  for measuring a surface temperature (outer-face temperature or outside-air temperature) of the wing leading edge  2   a  of the main wing  2 . 
     In addition, each of the bleed air supply line  45  and the heated air supply line  46  is provided with a check valve or the like (not illustrated) for preventing air from flowing back to the main engine  3  or the oil cooler (heat source)  26  from the hot air chamber  32  (hot air path  36 ). 
     Next, a description will be given of operation and effects of the above-configured aircraft ice protection system A of the present embodiment. 
     First, in the ice protection system A of the present embodiment, air S 1  is introduced from the air intake  31  during flight and fed to the oil cooler  26  by ram pressure, and heat exchange is performed in which the air S 1  is used as a cooling medium to cool hydraulic oil by the oil cooler  26 . At this time, the heat transfer pipe  30  of the oil cooler  26  is formed in a double pipe structure, by which the air S 1  flows through the flowing space  50  between the inner pipe  47  and the outer pipe  48 . Then, as described above, the low-temperature air S 1  passes through the flowing space  50  and flows along the inner pipe  47  through which the hydraulic oil kept at a high temperature flows. Thereby, heat exchange is efficiently performed between the hydraulic oil and the air S 1  to heat the air S 1 . 
     Further, in the present embodiment, the projected portions  49  are provided on the outer face of the inner pipe  47  by fixing the plurality of rings on the inner pipe  47 . Therefore, the inner pipe  47  is increased in heat transfer area, and the flow of the air S 1  which flows through the flowing space  50  between the inner pipe  47  and the outer pipe  48  is turned into a turbulent state, thereby facilitating heat exchange. Thereby, the heat exchange is performed at a higher efficiency, and the air S 1  is heated reliably and efficiently. 
     As described above, the air (hot air) S 1 ′ heated by the oil cooler  26  is supplied through the heated air supply line  46  to the other duct pipe  41  of the hot air chamber  32 , and injected into the hot air path  36  between the wing leading edge  2   a  and the guide plate  33  from the injection hole  41   a  of the other duct pipe  41  through the connecting pipe  43  and the opening  33   a  of the guide plate  33 . Then, the heated air S 1 ′ flows inside the hot air path  36 , by which the wing leading edge  2   a  is heated from the inside, thus making it possible to prevent ice accretion on the outside of the wing  2  or to remove the accreted ice. 
     As described above, in the ice protection system A of the present embodiment, the air S 1  which was conventionally discharged outside after being introduced from the air intake  31  and used as a cooling medium of the oil cooler  26  is utilized for ice protection. Thereby, when the air S 1 ′ heated by the oil cooler  26  can be used to carry out ice protection sufficiently, it is no longer necessary to use the bleed air S 2 . That is, when the thermometer  53  detects a temperature at which sufficient ice protection can be attained, the on-off valve  51  is closed and the on-off valve  52  is opened by control of the valve controller  54 . Therefore, only the air S 1 ′ is supplied to the hot air path  36  through the injection hole  41   a  of the other duct pipe  41 , the connecting pipe  43  and the opening  33   a  of the guide plate  33 . 
     On the other hand, when only the air S 1 ′ heated by the oil cooler  26  is not able to provide sufficient effects of ice protection, that is, when the thermometer  53  detects a temperature at which sufficient ice protection is not attained, the valve controller  54  controls the opening and closing of the on-off valves  51 ,  52  in response to the detection results. Thereby, the bleed air S 2  is injected to the hot air path  36  from the main engine  3  through the injection hole  40   a  of the one duct pipe  40 , the connecting pipe  42  and the opening  33   a  of the guide plate  33 . Then, since the bleed air S 2  is higher in temperature than the air S 1 ′ heated by the oil cooler  26 , it is possible to reliably carry out ice protection. 
     Therefore, in the aircraft ice protection system A of the present embodiment, as with a conventional case, the bleed air S 2  is supplied to the hot air chamber  32  from the main engine  3 , thus making it possible to protect the wing leading edge  2   a  from ice. Also, the air S 1  introduced from the air intake  31  installed on the airframe  28  is fed to the oil cooler  26  by ram pressure, thereby producing the hot air (heated air) S 1 ′. 
     Further, as described in the present embodiment where the oil cooler  26  in which the air S 1  introduced from the air intake  31  was conventionally used as a cooling medium and the air S 1 ′ after heat exchange was discharged outside is used as a heat source, heat produced in association with flight of the aircraft  1  is effectively used for ice protection. 
     Therefore, when the bleed air S 2  and the air S 1 ′ heated by the oil cooler  26  can be selectively supplied to the hot air chamber  32  to carry out ice protection. In addition, when the air S 1 ′ heated by the oil cooler  26  can be used to attain sufficient ice protection, it is no longer necessary to use the bleed air S 2 . Further, when only the air S 1 ′ heated by the oil cooler  26  is unable to attain sufficient ice protection, the bleed air S 2  is used to reliably carry out ice protection. Still further, even when the bleed air S 2  is used, the heated air S 1 ′ can be used to keep the wing  2  heat in advance (able to heat the wing  2  in advance), thereby realizing a state that ice accretion is less likely to take place. It is, thus, possible to suppress an extract amount of the bleed air S 2  to a minimum level. 
     According to the aircraft ice protection system A of the present embodiment and the aircraft  1  provided with the ice protection system A, the bleed air S 2  extracted from the main engine  3  becomes unnecessary or an extract amount of the bleed air S 2  is suppressed to a minimum necessary level during flight. Therefore, as compared with an aircraft provided with conventional ice protection equipment, it is possible to suppress a decrease in impelling force of the main engine and improve the performance of the aircraft such as a reduction in operation cost associated with an improvement in fuel consumption efficiency. 
     Further, in the aircraft ice protection system A of the present embodiment, the oil cooler  26  is configured in such a manner that the air S 1  introduced from the air intake  31  is allowed to flow between the outer pipe  48  and the inner pipe  47  of the double-pipe structured heat transfer pipe  30 . Thereby, heat exchange can be performed with hydraulic oil which flows through the inner pipe  47 , thereby reliably heating the air S 1 . 
     Still further, since the projected portions  49  are provided on the outer face of the inner pipe  47 , the inner pipe  47  is increased in heat transfer area. And, when the air S 1  flows between the outer pipe  48  and the inner pipe  47 , the projected portions  49  turn the flow into a turbulent state, thereby facilitating heat exchange. Accordingly, it is possible to increase the efficiency of heat exchange between the hydraulic oil flowing through the inner pipe  47  and the air S 1  flowing between the inner pipe  47  and the outer pipe  48 . Thus, the piping length inside the oil cooler  26  (heat transfer pipe length) can be decreased. Then, as described above, since the piping inside the oil cooler  26  is decreased in length, it is possible to reduce the weight of the airframe of the aircraft  1 . 
     Second Embodiment 
     Next, with reference to  FIG. 1 , and  FIG. 3  to  FIG. 7 , a description will be given of an aircraft ice protection system of a Second Embodiment in the present invention and an aircraft provided with the ice protection system. The aircraft ice protection system of the present embodiment is mainly different from the First Embodiment only in a heat source that heats air introduced from an air intake. Therefore, configurations similar to those of the First Embodiment will be given the same reference numerals, with a detailed description omitted here. 
     As shown in  FIG. 6 , the aircraft ice protection system B of the present embodiment selectively supplies bleed air S 2  extracted from a main engine  3  and air S 1 ′ introduced from an air intake  31  and heated at an oil tank (a heat source of an airframe of an aircraft)  27  to a hot air chamber  32  formed inside a wing leading edge  2   a  of a main wing  2 , thereby carrying out ice protection. 
     Further, as shown in  FIG. 7 , the oil tank  27  is provided with an outer shell  56  which forms a flowing space  50  between the outer shell  56  and an outer face of a tank main body  55  which pools hydraulic oil. Still further, the tank main body  55  is provided with projected portions (projected pieces)  57  on the outer face. 
     As shown in  FIG. 4  and  FIG. 6 , the other duct pipe  41  is connected to the oil tank  27  mounted on a hydraulic circuit  5  by a heated air supply line  46 . At this time, the other duct pipe  41  is connected to the flowing space  50  at which air S 1  between the tank main body  55  of the oil tank  27  and the outer shell  56  flows via the heated air supply line  46 . 
     In addition, one duct pipe  40  is connected to the main engine  3  by a bleed air supply line  45 , as with the First Embodiment. The bleed air S 2  of the main engine  3  is supplied from the one duct pipe  40  to a hot air path  36  (hot air chamber  32 ). 
     In the above-configured aircraft ice protection system B of the present embodiment, the air S 1  is introduced from the air intake  31  and fed into the oil tank  27  by ram pressure during flight. And, heat exchange is performed between the air S 1  and hydraulic oil which is temporarily pooled at the tank main body  55  of the oil tank  27 . At this time, the oil tank  27  is formed so as to have the tank main body  55  and the outer shell  56 , and the air S 1  flows through the flowing space  50  between the tank main body  55  and the outer shell  56 . Then, the air S 1  passes through the flowing space  50  and flows along the outer face of the tank main body  55  at which the hydraulic oil is pooled. Thus, heat is efficiently exchanged between the hydraulic oil and the air S 1  to heat the air S 1 . 
     Further, in the present embodiment, since the projected portions  57  are provided on the outer face of the tank main body  55 , the tank main body  55  is increased in heat transfer area. The flow of the air S 1  which flows through the flowing space  50  between the tank main body  55  and the outer shell  56  is turned into a turbulent state, thereby facilitating heat exchange. As a result, as with the First Embodiment, the heat exchange is performed at a higher efficiency to heat air reliably and efficiently. 
     As described above, the air (hot air) S 1 ′ heated at the oil tank  27  is supplied through the heated air supply line  46  to the other duct pipe  41  of the hot air chamber  32 , and injected into the hot air path  36  between the wing leading edge  2   a  and the guide plate  33  from an injection hole  41   a  of the other duct pipe  41  through a connecting pipe  43  and an opening  33   a  of the guide plate  33 . Then, the heated air S 1 ′ flows inside the hot air path  36 , by which the wing leading edge  2   a  is heated from inside to prevent ice accretion on the outside of the wing  2  or remove the accreted ice. 
     Thereby, in the ice protection system B of the present embodiment as well, when the air S 1 ′ heated at the oil tank  27  is able to carry out ice protection sufficiently, it is no longer necessary to use the bleed air S 2 . On the other hand, when only the air S 1 ′ heated at the oil tank  27  is not able to attain sufficient effects of ice protection, a valve controller  54  controls opening and closing of on-off valves  51 ,  52 , as with the First Embodiment. Thereby, the bleed air S 2  is injected into the hot air path  36  from the main engine  3  through an injection hole  40   a  of the one duct pipe  40 , the connecting pipe  42  and the opening  33   a  of the guide plate  33 . Since the bleed air S 2  is higher in temperature than the air S 1 ′ heated at the oil tank  27 , it is possible to reliably carry out ice protection. 
     Therefore, in the aircraft ice protection system B of the present embodiment, as with the First Embodiment, the bleed air S 2  is supplied from the main engine  3  to the hot air chamber  32 , thus making it possible to protect the wing leading edge  2   a  from ice. Further, the air S 1  introduced from the air intake  31  installed at the airframe  28  is fed into the oil tank  27  by ram pressure, thereby producing the hot air (heated air) S 1 ′ by utilizing the oil tank  27 . 
     Thereby, the bleed air S 2  and the air S 1 ′ heated at the oil tank  27  are selectively supplied to the hot air chamber  32 , thus making it possible to carry out ice protection. When the air S 1 ′ heated at the oil tank  27  can be used to carry out ice protection sufficiently, it is no longer necessary to use the bleed air S 2 . Further, when only the air S 1 ′ heated at the oil tank  27  is unable to attain sufficient ice protection, the bleed air S 2  is used to reliably carry out ice protection. Still further, even when the bleed air S 2  is used, the heated air S 1 ′ can be used to keep the wing  2  heat in advance, thereby realizing a state at which ice accretion is less likely to take place. Thus, it is possible to suppress an extract amount of the bleed air S 2  to a minimum level. 
     Therefore, in the aircraft ice protection system B of the present embodiment (and the aircraft  1  provided with the ice protection system B) as well, the bleed air S 2  extracted from the main engine  3  becomes unnecessary or an extract amount of the bleed air S 2  is suppressed to a necessary minimum level during flight. Consequently, as compared with an aircraft having conventional ice protection equipment, a decrease in impelling force of the main engine is suppressed to improve the performance of the aircraft such as a reduction in operation cost associated with an improvement in fuel consumption efficiency. 
     Further, as described in the present embodiment, when heat is exchanged with hydraulic oil which is pooled in the oil tank  27  to heat the air S 1 , the present invention performs not only cooling the hydraulic oil, which is at a high temperature, by an oil cooler  26  but also causing temperature decrease (making low in temperature) of the hydraulic oil inside the oil tank  27  by exchanging heat with the air S 1 . Thereby, the hydraulic oil can be delayed in oxidation and deterioration, and replacement interval of the hydraulic oil is prolonged. 
     Third Embodiment 
     Next, with reference to  FIG. 1 ,  FIG. 3  to  FIG. 5 ,  FIG. 8  and  FIG. 9 , a description will be given of an aircraft ice protection system of a Third Embodiment in the present invention and an aircraft provided with the ice protection system. In the aircraft ice protection system of the present embodiment, air introduced from an air intake is heated by an oil cooler and a main engine. That is, a heat source is different from the First Embodiment and Second Embodiment. Therefore, in the present embodiment as well, configurations similar to those of the First Embodiment and Second Embodiment will be given the same reference numerals, with a detailed description omitted here. 
     As shown in  FIG. 8 , the aircraft ice protection system C of the present embodiment selectively supplies bleed air S 2  extracted from a main engine  3  and air S 1 ′ introduced from an air intake  31  and heated by the main engine  3  and an oil cooler  26  (a heat source of an airframe of an aircraft) to a hot air chamber  32  formed inside a wing leading edge  2   a  of a main wing  2 , thereby carrying out ice protection. 
     Further, in the present embodiment, air S 1  introduced from the air intake  31  is branched and fed individually to the oil cooler  26  and the main engine  3  by ram pressure. The air S 1 ′ heated respectively by the oil cooler  26  and the main engine  3  is mixed and supplied to the other duct pipe  41  inside the hot air chamber  32 . 
     Then, the oil cooler  26  of the present embodiment is provided with a double-pipe structured heat transfer pipe  30 , as with the First Embodiment (refer to  FIG. 5 ). 
     On the other hand, in the main engine  3 , for example, a casing of a burner  11  is formed in a double-tubular structure. That is, as shown in  FIG. 9 , a tubular core engine casing  8  is provided with an outer shell  60  which forms a flowing space  50  between the outer shell  60  and an outer face of the core engine casing  8 . Further, a part having the outer shell  60  (the burner  11  in the present embodiment) is provided with first projected portions (projected portions)  61  such as fins which are arranged in a dispersed manner all over the outer face of the core engine casing  8 . Each of the first projected portions  61  is formed so as to project from the outer face of the core engine casing  8  to an inner face of the outer shell  60 . That is, the tip of the projected portion  61  is not in contact with the inner face of the outer shell  60 . 
     Further, in the present embodiment, in addition to the first projected portions  61 , a plurality of second projected portions (projected portions)  62  projecting from the outer face to the inner face of the outer shell  60  are installed on an outer face of the core engine casing  8  at the part having the outer shell  60 . Then, these second projected portions  62  are formed so as to extend in a circumferential direction at the center of an axis line O 1  of the core engine casing  8  and also gradually move from one end side to the other end side in the direction of the axis line O 1 . A predetermined clearance is also given between adjacent second projected portions  62 . Thereby, a helical flowing space  50  is formed between the second projected portions  62  which are adjacent with each other in the direction of the axis line O 1 . 
     Then, in the present embodiment, the other duct pipe  41  is connected to the oil cooler  26  and the main engine  3  by a heated air supply line  46 . At this time, the other duct pipe  41  is connected to a flowing space  50  between an inner pipe  47  and an outer pipe  48  of the oil cooler  26  and also to a flowing space  50  between the core engine casing  8  and the outer shell  60  of the main engine  3  via the heated air supply line  46 . 
     One duct pipe  40  is connected to the main engine  3  by the bleed air supply line  45 , as with the First and Second Embodiments. Bleed air S 2  of the main engine  3  is supplied from the one duct pipe  40  to the hot air path  36 . 
     In the above-configured aircraft ice protection system C of the present embodiment, the air S 1  is introduced from the air intake  31  during flight, branched and fed into the oil cooler  26  and the main engine  3  by ram pressure. The air S 1  is heated by exchanging heat with the oil cooler and heated also by exchanging heat with the core engine casing  8  of the main engine  3 . At this time, in the main engine  3 , the air S 1  flows through the plurality of flowing spaces  50  between the core engine casing  8  and the outer shell  60 . Then, since these flowing spaces  50  (the plurality of second projected portions  62 ) are formed in a helical manner, the air S 1  flowing through each of the flowing spaces  50  flows while swirling along the outer face of the tubular core engine casing  8 . Therefore, the air S 1  is efficiently heated by the outer face of the core engine casing  8  and the second projected portions  62 . 
     Further, in the present embodiment, since the first projected portions  61  are provided on the outer face of the core engine casing  8  of the main engine  3 , a heat transfer area is increased by the second projected portions  62  and the first projected portions  61 . Further, the flow of the air S 1  flowing through each of the flowing spaces  50  is turned into a turbulent state, thereby facilitating heat exchange. Thereby, the heat exchange is performed at a higher efficiency and the air is heated reliably and efficiently. 
     As described above, the air (hot air) S 1 ′ heated outside the core engine casing  8  of the main engine  3  is mixed with the air S 1 ′ heated by the oil cooler  26  and supplied to the other duct pipe  41  of the hot air chamber  32  through the heated air supply line  46 . Then, the air S 1 ′ is injected into the hot air path  36  between the wing leading edge  2   a  and the guide plate  33  from an injection hole  41   a  of the other duct pipe  41  through the connecting pipe  43  and an opening  33   a  of the guide plate  33 . The heated air S 1 ′ flows inside the hot air path  36 , by which the wing leading edge  2   a  is heated from inside, thus preventing ice accretion on the outside of the wing  2  or removing the accreted ice. 
     Thereby, in the ice protection system C of the present embodiment as well, when the air S 1 ′ heated by the main engine  3  (and the air S 1 ′ heated by the oil cooler  26 ) can be used to carry out ice protection sufficiently, it is no longer necessary to use the bleed air S 2 . On the other hand, when only the air S 1 ′ heated by the main engine  3  (and the air S 1 ′ heated by the oil cooler  26 ) is unable to attain sufficient effects of ice protection, as with the First Embodiment and Second Embodiment, the valve controller  54  controls opening and closing of the on-off valves  51 ,  52 . And, the bleed air S 2  is injected into the hot air path  36  between the wing leading edge  2   a  and the guide plate  33  from the main engine  3  through the injection hole  40   a  of one duct pipe  40 , the connecting pipe  42  and the opening  33   a  of the guide plate  33 . Since this bleed air S 2  is higher in temperature than the heated air S 1 ′, it is possible to reliably carry out ice protection. 
     In the present embodiment, the air S 1  is heated by flowing outside the core engine casing  8  of the main engine  3 . Therefore, as compared with a case where a heat source is made available only from the oil cooler  26  or the oil tank  27  on the hydraulic circuit  5  as found in the First Embodiment and Second Embodiment, the main engine  3  is designed so as not to cause a reduction in impelling force and the main engine  3  is used as a heat source. Thereby, the air S 1 ′ supplied to the hot air chamber  32  is made higher in temperature. As a result, effects of ice protection can be made greater than in the First Embodiment and Second Embodiment. 
     Therefore, in the aircraft ice protection system C of the present embodiment, as with the First Embodiment and Second Embodiment, the bleed air S 2  is supplied from the main engine  3  to the hot air chamber  32 , thus it makes possible to protect the wing leading edge  2   a  from ice. In addition, the air S 1  introduced from the air intake  31  installed on the airframe  28  is fed into the main engine  3  (and the oil cooler  26 ) by ram pressure, and the main engine  3  is utilized to produce the hot air (heated air) S 1 ′. 
     Thereby, the bleed air S 2  and the air S 1 ′ heated by the main engine  3  are selectively supplied to the hot air chamber  32  to carry out ice protection. And, when the air S 1 ′ heated by the main engine  3  is able to carry out ice protection sufficiently, it is no longer necessary to use the bleed air S 2 . Further, when only the air S 1 ′ heated by the main engine  3  is unable to carry out ice protection sufficiently, the bleed air S 2  can be used to reliably carry out ice protection. Still further, even when the bleed air S 2  is used, the heated air S 1 ′ can be used to keep the wing  2  heat in advance and realize a state in which ice accretion is less likely to take place. It is, therefore, possible to suppress an extract amount of the bleed air S 2  to a minimum level. 
     Therefore, in the aircraft ice protection system C of the present embodiment (and the aircraft  1  provided with the ice protection system C) as well, the bleed air S 2  extracted from main engine  3  is made unnecessary or an extract amount of the bleed air S 2  is suppressed to a minimum necessary level during flight. Therefore, as compared with an aircraft having conventional ice protection equipment, a decrease in impelling force of the main engine is suppressed to improve the performance of the aircraft such as a reduction in operation cost associated with an improvement in fuel consumption efficiency. 
     Further, as described in the present embodiment, the first projected portions  61  such as fins are provided on the outer face of the core engine casing  8  to increase a heat transfer area. Still further, when the air S 1  flows through the flowing space  50  between the core engine casing  8  and the outer shell  60 , the first projected portions  61  turn the flow into a turbulent state, thereby facilitating heat exchange. As a result, the heat exchange is performed at a higher efficiency with the air S 1  flowing through the flowing space  50 , thereby heating the air S 1  reliably and efficiently. Also, the heated air S 1 ′ is effectively used for ice protection. 
     Still further, the second projected portions  62  are formed in such a manner that the air S 1  is allowed to flow while swirling along the outer face of the core engine casing  8 . Therefore, contact time of the air S 1  on the core engine casing  8  is increased to raise the efficiency of heat exchange. And, the air S 1  is heated reliably and efficiently and the heated air S 1 ′ is effectively used for ice protection. 
     A description has been so far given of First, Second and Third Embodiments of the aircraft ice protection systems in the present invention. The present invention shall not be limited to the above-described First, Second and Third Embodiments but may be modified whenever necessary within a scope not departing from the gist of the present invention. 
     For example, as with the Third Embodiment when a heat source is provided by a combination of the main engine  3  and the oil cooler  26 , heat sources of First, Second and Third Embodiments are selectively combined. That is, it is acceptable that the air S 1  introduced from the air intake  31  be heated by using the oil cooler  26  and the oil tank  27 ; the oil tank  27  and the main engine  3 ; or the oil cooler  26 , the oil tank  27  and the main engine  3 ; and supplied to the hot air chamber  32 . In this case, it is also acceptable that the air S 1 ′ heated by a plurality of heat sources be individually supplied to the hot air chamber  32  (parallel system), or the air S 1  be heated sequentially by a plurality of heat sources and supplied to the hot air chamber  32  (serial system). 
     Further, in the First, Second and Third Embodiments, a description has been given of a case where the bleed air S 2  and the air S 1 ′ heated by a heat source are supplied to the hot air chamber  32  formed at the wing leading edge  2   a  of the main wing  2 , thereby protecting the main wing  2  from ice. As a matter of course, it is acceptable that the hot air chamber  32  be formed at a tail assembly or the like, and ice protection systems A, B, C similar to those of the present embodiments be used to supply the bleed air S 2  and the air S 1 ′ heated by a heat source to the hot air chamber  32  formed on the other wing (also to a hot air chamber  32  formed on the other wing), thereby attaining effects of ice protection. 
     Still further, in the Third Embodiment, the flowing space  50  for the air S 1  is formed so as to have the first projected portions  61  and the helical second projected portions  62 . Even in a case where projected portions relating to the present invention are provided as components, at least one of the first projected portions  61  and the second projected portions  62  are provided, thus making it possible to reliably and efficiently heat the air S 1 . In addition, the projected portions  49 ,  57 ,  61 ,  62  may not be necessarily formed in the flowing space  50 . 
     When the main engine  3  is used to heat the air S 1 , it is not always necessarily restricted that the burner  11  be used to heat the air S 1 . 
     DESCRIPTION OF REFERENCE SYMBOLS 
     
         
           1 : aircraft 
           2 : main wing (wing) 
           2   a : wing leading edge 
           3 : main engine (heat source)
         4 : actuator     5 : hydraulic circuit     
           5 ′: manifold 
           6 : gas turbine 
           7 : fan casing 
           8 : core engine casing 
           9 : fan 
           10 : compressor 
           11 : burner 
           12 : turbine 
           15 : rotating shaft 
           16 : fan blade 
           17 : low-pressure compressor 
           18 : high-pressure compressor 
           19 : high-pressure turbine 
           20 : low-pressure turbine 
           21 : first rotor shaft 
           22 : second rotor shaft 
           25 : hydraulic pump 
           26 : oil cooler (heat source) 
           27 : oil tank (heat source) 
           28 : body (airframe) 
           30 : heat transfer pipe 
           31 : air intake 
           32 : hot air chamber 
           33 : guide plate 
           33   a : opening 
           34 : partition wall 
           35 : partition wall 
           36 : hot air path 
           40 : duct pipe 
           40   a : injection hole 
           41 : duct pipe 
           41   a : injection hole 
           42 : connecting pipe 
           43 : connecting pipe 
           45 : bleed air supply line 
           46 : heated air supply line 
           47 : inner pipe 
           48 : outer pipe 
           49 : projected portion 
           50 : flowing space 
           51 : on-off valve 
           52 : on-off valve 
           53 : thermometer 
           54 : valve controller 
           55 : tank main body 
           56 : outer shell 
           60 : outer shell 
           61 : first projected portion (projected portion) 
           62 : second projected portion (projected portion) 
         A: aircraft ice protection system 
         B: aircraft ice protection system 
         C: aircraft ice protection system 
         S 1 : air 
         S 1 ′: heated air 
         S 2 : bleed air