Patent Application
20200248580
The present disclosure relates generally to gas turbine engines and, more particularly, to apparatus and methods used to detect duct ruptures within air management systems used with gas turbine engines.
Gas turbine engines, such as those used to power modern commercial and military aircraft, typically include a fan section, a compressor section, a combustor section and a turbine section. During operation, air is pressurized in the compressor section and mixed with fuel and burned in the combustor section to generate hot combustion gases. The hot combustion gases are then communicated through the turbine section, which extracts energy from the gases to power the compressor section, the fan section and various other systems occurring within or proximate the gas turbine engine. A typical engine core flow path extends sequentially through the compressor section, the combustor section and the turbine section. A bypass flow path is typically formed between a radial inner surface of a nacelle and a radial outer surface of a core engine case that contains the core air flow extending through the various sections. A fan within the fan section drives air through the bypass flow path.
Modern aircraft typically use the core air flow to provide pressurized air to other systems or components within an air management system. For example, pressurized air bled from one or more bleed ports mounted to the core engine case proximate the compressor section may be routed via conduits or ducts to provide environmental air within the cabin of the aircraft or anti-icing air to a nose-lip section of the nacelle surrounding the core engine case and fan. Oftentimes, the pressurized air bled from the one or more bleed ports is at an elevated temperature, which may approach a fire detection threshold temperature. Accordingly, a rupture within a duct or valve in the proximity of the core engine case or within the confines of the nacelle may inadvertently trigger a fire detection system, causing a needless and potentially costly release of an extinguishment or an inflight shutdown (IFSD).
On the other hand, if the temperature of the pressurized gas is sufficiently below the fire detection threshold temperature, the ruptured duct or valve may go undetected, also leading to costly repairs to an engine or to an airframe component proximate the engine in the event a part breaks free from the engine. Further complicating matters is the presence of overpressure release mechanisms configured to exhaust the pressurized air from a ruptured duct away from the nacelle, rendering overpressure detectors ineffective. Accordingly, a temperature detection system tuned to detect a ruptured duct or valve in the proximity of a core engine case may be used to detect the rupture and also to avoid a possible false fire detection and subsequent release of extinguishment or an IFSD.
A system for detecting a ruptured duct transporting a high-temperature fluid within a gas turbine engine is disclosed. In various embodiments, the system includes a rupture detection line configured to extend within a first fire zone of the gas turbine engine; a plurality of rupture sensing elements in electrical communication with and disposed along the rupture detection line, the plurality of rupture sensing elements configured to detect a presence of a heated fluid having a heated fluid temperature less than a fire temperature; and a processor configured to monitor the plurality of rupture sensing elements.
In various embodiments, the rupture detection line is configured to extend into a second fire zone. In various embodiments, the rupture detection line includes a first section configured to detect a rupture of a first duct in the first fire zone and a second section configured to detect the rupture of a second duct in the second fire zone. In various embodiments, the rupture detection line includes a first section configured to detect a rupture of a first duct in the first fire zone and a second section configured to detect a flow of the heated fluid from the first duct into the second fire zone. In various embodiments, the first ruptured duct is one of a cooling air duct, an environmental air duct a de-icing duct and an oil duct. In various embodiments, the second ruptured duct is one of an environmental air duct and a de-icing duct.
In various embodiments, the rupture detection line includes a first section within the first fire zone and a second section within a second fire zone. In various embodiments, the plurality of rupture sensing elements includes a first plurality of rupture sensing elements disposed on the first section and a second plurality of rupture sensing elements disposed on the second section. In various embodiments, the first fire zone is a core compartment and the second fire zone is a pylon. In various embodiments, the first plurality of sensing elements includes a first thermistor configured to detect a first temperature and the second plurality of sensing elements includes a second thermistor configured to detect a second temperature. In various embodiments, the first temperature is characteristic of a heated gas bled from a high pressure compressor of the gas turbine engine.
A gas turbine engine is disclosed. In various embodiments, the gas turbine engine includes a compressor section; a duct configured to bleed air from the compressor section; a rupture detection line configured to extend within a first fire zone of the gas turbine engine; a plurality of rupture sensing elements in electrical communication with and disposed along the rupture detection line, the plurality of rupture sensing elements configured to detect a presence of a heated fluid escaping from the duct, the heated fluid having a heated fluid temperature less than a fire temperature; and a processor configured to monitor the plurality of rupture sensing elements.
In various embodiments, the rupture detection line is configured to extend into a second fire zone. In various embodiments, the first fire zone is a core compartment and the second fire zone is a pylon. In various embodiments, the rupture detection line includes a first section within the first fire zone and a second section within the second fire zone and wherein the plurality of rupture sensing elements includes a first plurality of rupture sensing elements disposed on the first section and a second plurality of rupture sensing elements disposed on the second section. In various embodiments, the first plurality of sensing elements includes a first thermistor configured to detect a first temperature and the second plurality of sensing elements includes a second thermistor configured to detect a second temperature. In various embodiments, the first temperature is characteristic of a heated gas bled from a high pressure compressor of the gas turbine engine.
A thermal detection system for an aircraft is disclosed. In various embodiments, the thermal detection system includes a fire detection system having a fire detection line disposed within a first fire zone of a gas turbine engine, a fire sensing element, configured to detect a fire occurring within the first fire zone and a first processor configured to monitor the fire sensing element; and a duct rupture detection system having a rupture detection line disposed within the first fire zone, a rupture sensing element in electrical communication with and disposed along the rupture detection line, the rupture sensing element configured to detect a presence of a heated fluid having a heated fluid temperature less than a fire temperature, and a second processor configured to monitor the rupture sensing element. In various embodiments, the rupture detection line includes a first section configured to detect a ruptured duct in the first fire zone and a second section configured to detect a flow of the heated fluid from the ruptured duct into a second fire zone. In various embodiments, the rupture detection line is monitored by an electronic engine control or a full authority digital engine control.
The subject matter of the present disclosure is particularly pointed out and distinctly claimed in the concluding portion of the specification. A more complete understanding of the present disclosure, however, may best be obtained by referring to the following detailed description and claims in connection with the following drawings. While the drawings illustrate various embodiments employing the principles described herein, the drawings do not limit the scope of the claims.
The following detailed description of various embodiments herein makes reference to the accompanying drawings, which show various embodiments by way of illustration. While these various embodiments are described in sufficient detail to enable those skilled in the art to practice the disclosure, it should be understood that other embodiments may be realized and that changes may be made without departing from the scope of the disclosure. Thus, the detailed description herein is presented for purposes of illustration only and not of limitation. Furthermore, any reference to singular includes plural embodiments, and any reference to more than one component or step may include a singular embodiment or step. Also, any reference to attached, fixed, connected, or the like may include permanent, removable, temporary, partial, full or any other possible attachment option. Additionally, any reference to without contact (or similar phrases) may also include reduced contact or minimal contact. It should also be understood that unless specifically stated otherwise, references to “a,” “an” or “the” may include one or more than one and that reference to an item in the singular may also include the item in the plural. Further, all ranges may include upper and lower values and all ranges and ratio limits disclosed herein may be combined.
Referring now to the drawings,
The gas turbine engine 20 generally includes a low speed spool 30 and a high speed spool 32 mounted for rotation about an engine central longitudinal axis A relative to an engine static structure 36 via several bearing systems 38. It should be understood that various bearing systems at various locations may alternatively or additionally be provided and the location of the several bearing systems 38 may be varied as appropriate to the application. The 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. The inner shaft 40 is connected to the fan 42 through a speed change mechanism, which in this gas turbine engine 20 is illustrated as a fan drive gear system 48 configured to drive the fan 42 at a lower speed than the low speed spool 30. The high speed spool 32 includes an outer shaft 50 that interconnects a high pressure compressor 52 and a high pressure turbine 54. A combustor 56 is arranged in the gas turbine engine 20 between the high pressure compressor 52 and the high pressure turbine 54. A mid-turbine frame 57 of the engine static structure 36 is arranged generally between the high pressure turbine 54 and the low pressure turbine 46 and may include airfoils 59 in the core flow path C for guiding the flow into the low pressure turbine 46. The mid-turbine frame 57 further supports the several bearing systems 38 in the turbine section 28. The inner shaft 40 and the outer shaft 50 are concentric and rotate via the several bearing systems 38 about the engine central longitudinal axis A, which is collinear with longitudinal axes of the inner shaft 40 and the outer shaft 50.
The air in the core flow path C is compressed by the low pressure compressor 44 and then the high pressure compressor 52, mixed and burned with fuel in the combustor 56, and then expanded over the high pressure turbine 54 and low pressure turbine 46. The low pressure turbine 46 and the high pressure turbine 54 rotationally drive the respective low speed spool 30 and the high speed spool 32 in response to the expansion. It will be appreciated that each of the positions of the fan section 22, the compressor section 24, the combustor section 26, the turbine section 28, and the fan drive gear system 48 may be varied. For example, the fan drive gear system 48 may be located aft of the combustor section 26 or even aft of the turbine section 28, and the fan section 22 may be positioned forward or aft of the location of the fan drive gear system 48.
Referring now to
Various components may be provided in the core compartment 210, such as, for example, a cooling air duct 212, used to route pressurized air from the compressor section to the turbine section, or an environmental air duct 214, used to route pressurized air from the compressor section, through the pylon 206, and ultimately to the cabin of an aircraft. The cooling air duct 212 is under high pressure and may supply compressed air, for example, from a low pressure compressor or a high pressure compressor to a high pressure turbine for cooling. Similarly, the environmental air duct 214 is under high pressure and may supply compressed air, for example, from the compressor section of the core engine 204 to an Environmental Control System (ECS) for the aircraft. Other examples of pressurized air may be routed by conduit to a de-icing or anti-icing system (collectively referred to herein as a de-icing system including a de-icing duct for routing the pressurized air). A bypass flow path B, such as, for example, the bypass flow path B, described above with reference to
Still referring to
In various embodiments, the fan duct inner structure 208 that defines Zone A may include an upper bifurcation 232 and a lower bifurcation 234, each of which extend in a generally radial direction within the bypass flow path B and are configured to accommodate wires, conduits, engine mountings or other components. In various embodiments, the upper bifurcation 232 and the lower bifurcation 234 are configured to carry conduits transporting pressurized air or flammable fluids from on fire zone to another. For example, in various embodiments, the upper bifurcation 232 may be configured to route conduits—e.g., the environmental air duct 214—from positions within the first fire zone 222 (e.g., Zone A) to positions within the second fire zone (e.g., Zone B). The upper bifurcation 232 and the lower bifurcation 234 also facilitate opening of the nacelle 202 to access the core engine 204 or the various systems housed within the nacelle 202. In various embodiments, fire seals may also be positioned between the fire zones to hinder the spread of fire or the leaking of pressurized or flammable fluids from a ruptured duct from one zone to another.
Referring now to
In various embodiments, the one or more rupture sensing elements 358 may be disposed throughout an applicable fire zone and mounted to appropriate structure. For example, the one or more rupture sensing elements 358 may be disposed throughout the first fire zone 322 and mounted to a fan duct inner structure 308, such as, for example, the fan duct inner structure 208 described above with reference to
In various embodiments, the rupture detection system 350 may be a component within a thermal detection system 370, which compliments the rupture detection system 350 and includes separate lines dedicated to fire detection. For example, in various embodiments, the thermal detection system 370 may include a fire detection line 372. Spaced along the fire detection line 372 is one or more fire sensing elements 374, which may comprise a thermistor 376 (or a fire sensing element), similar to the thermistor 360 described above, but having different thermal properties. For example, the thermistor 360 coupled to the rupture detection line 352 may possess thermal properties that exhibit a substantial reduction in electrical resistance at a heated gas temperature characteristic of a ruptured duct (e.g., 1,000° F. or 537° C.), while the thermistor 376 coupled to the fire detection line 372 may possess thermal properties that exhibit a reduction in electrical resistance at a fire temperature characteristic of a fire (e.g., 2,000° F. or 1,093° C.).
In various embodiments, the rupture detection line 352 is connected to a rupture detection processor 362. The rupture detection processor 362 may comprise a full authority digital engine control (FADEC) system or an electronic engine control (EEC) system. Similarly, in various embodiments, the fire detection line 372 is connected to a fire detection processor 364. The fire detection processor 364 may be a part of or incorporated within the rupture detection processor 362 or it may comprise a second processor, separate and apart from the rupture detection processor 362. For example, in various embodiments, the rupture detection processor 362 may be a component of a gas turbine engine, while the fire detection processor 364 may be a component of the airframe upon which the gas turbine engine is mounted. Employing separate processors facilitates the rupture detection system 350 to be incorporated into a previously existing fire detection system, having a capability to detect fire within a nacelle structure of a gas turbine engine but no capability to detect hot gases being expelled from a ruptured duct that often times exhibit a characteristic temperature less than the characteristic temperature of a fire.
Referring now to
A rupture detection processor 462, similar to the rupture detection processor 362 described above with reference to
The rupture detection processor 462 is also configured to monitor the one or more rupture sensing elements 458 disposed in the second fire zone 428. In the event a ruptured duct situation occurs within the first fire zone 422 (e.g., the core compartment 410), the heated and pressurized air may breach a fire seal disposed between the two fire zones, thereby facilitating, in various embodiments, the heated and pressurized air to flow into the second fire zone 428 (e.g., the pylon 406). In such event, a second zone rupture alarm signal—e.g., a second light or illuminator 482—is provided to the cockpit and activated to alert the cockpit crew. In the event that one or both of the first fire zone 422 and the second fire zone 428 indicate a ruptured duct situation, appropriate action may be taken, including shutting down the engine or throttling back the speed of the engine to idle. In various situations, such action may avoid inadvertent or unnecessary release of extinguishment and the associated expenses involved therewith. Additionally, in various embodiments, an airframe controller 463, separate from the rupture detection processor 462, may be configured to monitor the one or more rupture sensing elements 458 disposed in the second fire zone 428. In such embodiments, rupture detection within an engine is monitored by an engine controller (or first processor) and rupture detection within a pylon is monitored by an airframe controller.
Benefits, other advantages, and solutions to problems have been described herein with regard to specific embodiments. Furthermore, the connecting lines shown in the various figures contained herein are intended to represent exemplary functional relationships and/or physical couplings between the various elements. It should be noted that many alternative or additional functional relationships or physical connections may be present in a practical system. However, the benefits, advantages, solutions to problems, and any elements that may cause any benefit, advantage, or solution to occur or become more pronounced are not to be construed as critical, required, or essential features or elements of the disclosure. The scope of the disclosure is accordingly to be limited by nothing other than the appended claims, in which reference to an element in the singular is not intended to mean “one and only one” unless explicitly so stated, but rather “one or more.” Moreover, where a phrase similar to “at least one of A, B, or C” is used in the claims, it is intended that the phrase be interpreted to mean that A alone may be present in an embodiment, B alone may be present in an embodiment, C alone may be present in an embodiment, or that any combination of the elements A, B and C may be present in a single embodiment; for example, A and B, A and C, B and C, or A and B and C. Different cross-hatching is used throughout the figures to denote different parts but not necessarily to denote the same or different materials.
Systems, methods and apparatus are provided herein. In the detailed description herein, references to “one embodiment,” “an embodiment,” “various embodiments,” etc., indicate that the embodiment described may include a particular feature, structure, or characteristic, but every embodiment may not necessarily include the particular feature, structure, or characteristic. Moreover, such phrases are not necessarily referring to the same embodiment. Further, when a particular feature, structure, or characteristic is described in connection with an embodiment, it is submitted that it is within the knowledge of one skilled in the art to affect such feature, structure, or characteristic in connection with other embodiments whether or not explicitly described. After reading the description, it will be apparent to one skilled in the relevant art(s) how to implement the disclosure in alternative embodiments.
Furthermore, no element, component, or method step in the present disclosure is intended to be dedicated to the public regardless of whether the element, component, or method step is explicitly recited in the claims. No claim element herein is to be construed under the provisions of 35 U.S.C. 112(f) unless the element is expressly recited using the phrase “means for.” As used herein, the terms “comprises,” “comprising,” or any other variation thereof, are intended to cover a non-exclusive inclusion, such that a process, method, article, or apparatus that comprises a list of elements does not include only those elements but may include other elements not expressly listed or inherent to such process, method, article, or apparatus.
Finally, it should be understood that any of the above described concepts can be used alone or in combination with any or all of the other above described concepts. Although various embodiments have been disclosed and described, one of ordinary skill in this art would recognize that certain modifications would come within the scope of this disclosure. Accordingly, the description is not intended to be exhaustive or to limit the principles described or illustrated herein to any precise form. Many modifications and variations are possible in light of the above teaching.
