Aircraft compartment overheating indicator, system, and method

Abstract

The aircraft compartment is equipped with a self-generating thermal indicator that is activated when exposed to or otherwise encountered ambient temperatures exceeding a selected threshold temperature, and is visible in at least one area from and outside the aircraft. [Solution] A single-use, replaceable, self-generating aircraft compartment thermal indicator provides a localized, visually detectable signal that can be perceived by personnel outside the aircraft's components, is activated to signal on demand the presence of temperatures in the aircraft compartment exceeding a threshold temperature, and can be tested in situ.

Description

Technical Field

[0001] Government Rights This invention was made with government support under contract (FA8628-19-D-1000-FA810722F0001) awarded by the United States Department of Defense. The government has certain rights in this invention.

[0002]

[0001] This disclosure relates generally to the field of temperature indication in aircraft compartments. In particular, this disclosure relates to the field of temperature indication in specific aircraft compartments for indicating excessive temperature in aircraft compartments.

Background Art

[0003]

[0002] The ambient and operating temperatures present within an aircraft can be monitored within the cabin compartments, storage compartments, cargo compartments, and machinery compartments to promote safety and ensure that the temperature within the monitored compartments is maintained within an acceptable temperature range.

[0004]

[0003] System feedback from a temperature monitoring system can be relayed to aircraft personnel located within the cockpit. In an aircraft or other vehicle type, the gross weight is a factor to be considered, and the in-flight power usage and consumption to operate an electric heat monitoring system, as well as the weight of the electric monitoring system itself (monitoring system hardware, linkages, connections, wiring, etc.) can create additional weight that is undesirable for an aircraft and can significantly increase operating costs. Unless expressly stated as prior art, the description herein is not to be regarded as prior art merely by virtue of its inclusion in the technical field and / or background art.

Summary of the Invention

[0005]

[0004] Multiple embodiments of this invention relate to a self-generating thermal indicator in an aircraft compartment that can be activated when a region of the thermal indicator is exposed to or otherwise encountered an ambient temperature exceeding a selected threshold temperature, which may include a threshold temperature range. In this case, at least a region of the thermal indicator is visible from and outside the aircraft.

[0006]

[0005] This embodiment relates to a self-generating aircraft compartment thermal indicator (20) for identifying a temperature exceeding a threshold in an internal region of an aircraft. In this case, the aircraft compartment thermal indicator comprises an indicator body having an indicator body length ("l"). The indicator body further comprises an indicator body interior, an indicator body exterior, a first end of the indicator body (e.g., a display side), a second end of the indicator body, a second end base of the indicator body, and an internal piston. The internal piston further comprises a piston base having a first side and a second side of the piston base, and the internal piston further comprises a piston rod having a first end and a second end of the piston rod, the second end of the piston rod being at least connected to the first side of the piston base. The indicator body further comprises an indicator body wall, which is an indicator body wall extending along the length of the indicator body from a first end of the indicator body to a second end of the indicator body, and having an indicator body wall outer surface on the outside of the indicator body and an indicator body wall inner surface on the inside of the indicator body. The indicator body further comprises a battery compartment, which comprises a battery compartment base, and is defined by the battery compartment base and the indicator body wall inner surface. The indicator body further comprises a first chamber of the indicator body, which is defined by the indicator body wall inner surface, a first side of the piston base, and the battery compartment base, and comprises a fragile container for holding a certain amount of battery electrolyte. The indicator body further comprises a second chamber of the indicator body, which is defined by the indicator body wall inner surface and a second side of the piston base, and comprises a spring, which comprises a first end of the spring fixedly attached to the piston base, and a second end of the spring positioned directly adjacent to the second end base of the indicator body.A second chamber of the indicator body comprises a disassemblable retainer, which is connected to a spring and configured to hold the spring in a compressed spring state within a first temperature range, and to disassemble at a threshold temperature within a second temperature range, further configured to hold the spring in a compressed spring state while within the first temperature range, and further configured to release the spring from a compressed spring state to an extended spring state at a threshold temperature within the second temperature range, wherein the threshold temperature is higher than the first temperature range.

[0007]

[0006] In another embodiment, the aircraft compartment includes an interior area of ​​the aircraft, which includes a localized area within the aircraft compartment.

[0008]

[0007] In another embodiment, the thermal indicator of the aircraft compartment is self-generating.

[0009]

[0008] In another embodiment, the thermal indicator of the aircraft compartment further comprises a luminaire that is connected to the first end of the piston rod.

[0010]

[0009] In another embodiment, the lighting fixture is a light-emitting diode (LED).

[0011]

[0010] In another embodiment, the battery compartment further comprises a plurality of battery plates.

[0012]

[0011] In another embodiment, the decomposable retainer comprises at least one of a low-temperature alloy and a wax, and the at least one of the low-temperature alloy and the wax and the decomposable retainer have a melting point in the range of about 140 degrees Fahrenheit to about 165 degrees Fahrenheit.

[0013]

[0012] In another embodiment, the cryogenic alloy includes a cryogenic metal alloy comprising at least one of bismuth, lead, tin, indium, cadmium, thallium, gallium, and combinations thereof, wherein the cryogenic alloy has a melting point in the range of about 140 degrees Fahrenheit to about 165 degrees Fahrenheit.

[0014]

[0013] In another embodiment, the low-temperature alloy includes at least one of Rose alloy, Cerosafe, Wood alloy, Fields alloy, Cellolow 136, Cellolow 117, Gallium, and combinations thereof.

[0015]

[0014] In a further embodiment, the battery plate in the presence of the electrolyte (34a) is configured to form a battery configured to supply current.

[0016]

[0015] In another embodiment, the battery is a fuse battery.

[0017]

[0016] In another embodiment, the thermal indicator of the aircraft compartment further comprises an integrated thermal induction tube in direct contact with a disassemblable retainer.

[0018]

[0017] In another embodiment, the thermal indicator of the aircraft compartment is detachably attached to the aircraft's substrate.

[0019]

[0018] In a further embodiment, the thermal indicator of the aircraft compartment may be tested and inspected in situ.

[0020]

[0019] In another embodiment, the thermal indicator of the aircraft compartment further comprises an integrated thermal induction tube having a first end of a thermal induction tube that is fixedly extending into a second chamber of the indicator body, and a second end of a thermal induction tube that extends a selected distance from the second end base of the indicator body.

[0021]

[0020] In another embodiment, the thermal induction tube further comprises a thermal induction tube wall which includes a thermal induction tube wall outer periphery that is substantially adjacent to the disassemblable retainer in the second chamber of the indicator body.

[0022]

[0021] Further embodiments of this invention relate to an aircraft component comprising a self-generating aircraft compartment thermal indicator for identifying a temperature exceeding a threshold in an internal region of the aircraft. In this case, the aircraft compartment thermal indicator comprises an indicator body having an indicator body length ("l"). The indicator body further comprises an interior of the indicator body, an exterior of the indicator body, a first end of the indicator body (e.g., a display side), a second end of the indicator body, and a base of the second end of the indicator body, and an internal piston. The internal piston further comprises a piston base, which comprises a piston base having a first side and a second side of the piston base, and the internal piston further comprises a piston rod, which comprises a piston rod having a first end and a second end of the piston rod, the second end of the piston rod being at least connected to the first side of the piston base. The indicator body further comprises an indicator body wall, which is an indicator body wall extending along the length of the indicator body from a first end of the indicator body to a second end of the indicator body, and having an indicator body wall outer surface on the outside of the indicator body and an indicator body wall inner surface on the inside of the indicator body. The indicator body further comprises a battery compartment, which comprises a battery compartment base, and is defined by the battery compartment base and the indicator body wall inner surface. The indicator body further comprises a first chamber of the indicator body, which is defined by the indicator body wall inner surface, a first side of the piston base, and the battery compartment base, and comprises a fragile container for holding a certain amount of battery electrolyte.The indicator body further comprises a second chamber of the indicator body defined by the inner surface of the indicator body wall and a second side of the piston base, the second chamber of the indicator body comprising a spring having a first end fixedly attached to the piston base, and a second end of the spring positioned directly adjacent to the base of the second end of the indicator body. The second chamber of the indicator body further comprises a disassemblable retainer communicating with the spring and configured to hold the spring in a compressed spring state within a first temperature range, and to disassemble at a threshold temperature within a second temperature range, further configured to hold the spring in a compressed spring state while within the first temperature range, and further configured to release the spring from a compressed spring state to an expanded spring state at a threshold temperature within the second temperature range, wherein the threshold temperature is higher than the first temperature range.

[0023]

[0022] In another embodiment, the component of the aircraft is a component of the aircraft's wing.

[0024]

[0023] Another embodiment of this invention relates to an aircraft equipped with a self-generating aircraft compartment thermal indicator for identifying a temperature exceeding a threshold in an internal region of the aircraft. In this case, the aircraft compartment thermal indicator comprises an indicator body having an indicator body length ("l"). The indicator body further comprises an indicator body interior, an indicator body exterior, a first end of the indicator body (e.g., a display side), a second end of the indicator body, and a second end base of the indicator body, and an internal piston. The internal piston further comprises a piston base, which comprises a piston base having a first side and a second side of the piston base, and the internal piston further comprises a piston rod, which comprises a piston rod having a first end and a second end of the piston rod, the second end of the piston rod being at least connected to the first side of the piston base. The indicator body further comprises an indicator body wall, which is an indicator body wall extending along the length of the indicator body from a first end of the indicator body to a second end of the indicator body, and having an indicator body wall outer surface on the outside of the indicator body and an indicator body wall inner surface on the inside of the indicator body. The indicator body further comprises a battery compartment, which comprises a battery compartment base, and is defined by the battery compartment base and the indicator body wall inner surface. The indicator body further comprises a first chamber of the indicator body, which is defined by the indicator body wall inner surface, a first side of the piston base, and the battery compartment base, and comprises a fragile container for holding a certain amount of battery electrolyte. The indicator body further comprises a second chamber of the indicator body, which is defined by the indicator body wall inner surface and a second side of the piston base, and comprises a spring, which comprises a first end of the spring fixedly attached to the piston base, and a second end of the spring positioned directly adjacent to the second end base of the indicator body.A second chamber of the indicator body comprises a disassemblable retainer, which is connected to a spring and configured to hold the spring in a compressed spring state within a first temperature range, and to disassemble at a threshold temperature within a second temperature range, further configured to hold the spring in a compressed spring state while within the first temperature range, and further configured to release the spring from a compressed spring state to an extended spring state at a threshold temperature within the second temperature range, wherein the threshold temperature is higher than the first temperature range.

[0025]

[0024] Another embodiment of this invention relates to an aircraft compartment thermal indicator system (hereinafter referred to herein as "aircraft compartment temperature indicator system") for visually detecting, from the outside of the aircraft, a temperature exceeding a threshold in an aircraft compartment adjacent to the outside of the aircraft. The aircraft compartment temperature indicator system comprises a self-generating aircraft compartment thermal indicator mounted on an aircraft component, configured to identify a temperature exceeding a threshold in an internal compartment area of ​​the aircraft, and comprising an indicator body having an indicator body length ("l"). The indicator body further comprises an indicator body interior, an indicator body exterior, a first end of the indicator body, a second end of the indicator body, and a second end base of the indicator body, and an internal piston. The internal piston further comprises a piston base comprising a piston base having a first side and a second side of the piston base, and the internal piston further comprises a piston rod comprising a piston rod having a first end and a second end of the piston rod, the second end of the piston rod being at least connected to the first side of the piston base. The indicator body further comprises an indicator body wall, which is an indicator body wall extending along the length of the indicator body from a first end of the indicator body to a second end of the indicator body, and having an indicator body wall outer surface on the outside of the indicator body and an indicator body wall inner surface on the inside of the indicator body. The indicator body further comprises a battery compartment, which comprises a battery compartment base, and is defined by the battery compartment base and the indicator body wall inner surface. The indicator body further comprises a first chamber of the indicator body, which is defined by the indicator body wall inner surface, a first side of the piston base, and the battery compartment base, and comprises a fragile container for holding a certain amount of battery electrolyte.The indicator body further includes a second chamber of the indicator body defined by the inner surface of the indicator body wall and the second side of the piston base, the second chamber including a spring having a first end fixedly attached to the piston base and a second end of the spring disposed directly adjacent to the second end base of the indicator body. The second chamber of the indicator body further includes a frangible retainer that communicates with the spring and is configured to hold the spring in a compressed spring state within a first temperature range and to break down at a threshold temperature within a second temperature range, and is further configured to hold the spring in a compressed spring state while in the first temperature range and to release the spring from the compressed spring state toward an extended spring state at the threshold temperature within the second temperature range. The system further includes an integrated visually detectable signal device disposed within the first end of the indicator body, in communication with an electrical circuit, further disposed external to the aircraft, and configured to activate at a temperature above a threshold temperature, the threshold temperature being higher than the first temperature range.

[0026]

[0025] In another aspect, the battery compartment includes a plurality of battery plates.

[0027]

[0026] In another aspect, the visually detectable signal device includes a lighting fixture.

[0028]

[0027] In another aspect, the visually detectable signal device includes a light emitting diode (LED).

[0029]

[0028] In a further aspect, the signal device is configured to emit a light ray in response to the second end of the indicator being exposed to a threshold temperature in the range of approximately 140 degrees Fahrenheit to approximately 165 degrees Fahrenheit.

[0030]

[0029] In another aspect, the thermal indicator of the aircraft compartment is detachably attached to a component of the aircraft.

[0031]

[0030] In another main aspect, a thermal indicator of a self-powered compartment is configured to be removed from an aircraft component from a position outside the aircraft.

[0032]

[0031] In another main aspect, a thermal indicator of a self-powered aircraft compartment is accessible for inspection outside the aircraft.

[0033]

[0032] In another main aspect, the battery is a fuse battery, and the fuse battery is configured to activate at a threshold temperature.

[0034]

[0033] Another embodiment relates to a method for detecting a threshold temperature outside the aircraft in an internal region of an aircraft component which may include an aircraft compartment located in close proximity to the outside of the aircraft. In this case, the method includes placing a self-generating aircraft compartment thermal indicator inside the aircraft component, the aircraft component including the outer surface of the aircraft component (e.g., the outer surface of the aircraft), the outer surface of the aircraft component being located adjacent to an internal region of the aircraft component which may include an aircraft compartment, and the aircraft compartment thermal indicator comprises an indicator body having an indicator body length ("l"). The indicator body further comprises an indicator body interior, an indicator body exterior, a first end of the indicator body (e.g., the display side), a second end of the indicator body, and a second end base of the indicator body, and an internal piston. The internal piston further comprises a piston base having a first side and a second side of the piston base, and the internal piston further comprises a piston rod having a first end and a second end of the piston rod, the second end of the piston rod being at least connected to the first side of the piston base. The indicator body further comprises an indicator body wall having an indicator body wall extending along the length of the indicator body from the first end to the second end of the indicator body, and having an indicator body wall outer surface on the outside of the indicator body and an indicator body wall inner surface on the inside of the indicator body. The indicator body further comprises a battery compartment having a battery compartment base, and a battery compartment defined by the battery compartment base and the indicator body wall inner surface. The indicator body further comprises a first chamber of the indicator body having a fragile container for holding a certain amount of battery electrolyte, defined by the indicator body wall inner surface, the first side of the piston base, and the battery compartment base.The indicator body further comprises a second chamber of the indicator body defined by the inner surface of the indicator body wall and a second side of the piston base, the second chamber of the indicator body comprising a spring having a first end fixedly attached to the piston base, and a second end of the spring positioned directly adjacent to the base of the second end of the indicator body. The second chamber of the indicator body further comprises a disassemblable retainer communicating with the spring and configured to hold the spring in a compressed spring state within a first temperature range, and to disassemble at a threshold temperature within a second temperature range, further configured to hold the spring in a compressed spring state while within the first temperature range, and further configured to release the spring from a compressed spring state to an expanded spring state at a threshold temperature within the second temperature range, wherein the threshold temperature is higher than the first temperature range. The method further includes: disassembling a disassemblable retainer in the presence of a compartment temperature exceeding a threshold temperature; releasing a spring from a compressed spring state to an extended spring state; advancing a piston from an initial piston position to a piston deployed position; rupturing a fragile container to form a ruptured fragile container; releasing at least a portion of a certain amount of battery electrolyte from the ruptured fragile container; introducing at least a portion of a certain amount of battery electrolyte into a battery compartment to form a battery, the battery being configured to generate and supply current; leading a first end of a piston rod through a through-opening in the first end of an indicator body to a selected distance beyond the outer surface of an aircraft component; and introducing current from the battery to a light-emitting device to emit light from the light-emitting device.

[0035]

[0034] In another present embodiment, the method further includes visually detecting light emitted from a light-emitting device outside the aircraft.

[0036]

[0035] In another present embodiment, the method further includes detachably arranging a self-generating aircraft compartment thermal indicator within an aircraft component.

[0037]

[0036] In another embodiment, the method involves visually detecting illumination light from an aircraft compartment thermal indicator outside the aircraft to prove that the temperature inside the aircraft's components has exceeded at least a threshold temperature.

[0038]

[0037] The above-described features, functions, and advantages can be realized individually in various embodiments or in combination in yet another set of embodiments, and these details can be found by referring to the following description and accompanying drawings.

[0039]

[0038] Modifications of the present disclosure have been described in general terms, but now refer to the attached drawings. These are not necessarily drawn to scale. [Brief explanation of the drawing]

[0040] [Figure 1]

[0039] This is a diagram of a vehicle taking the form of an aircraft according to this embodiment. [Figure 2]

[0040] This is a cross-sectional side view of a self-generating aircraft compartment thermal indicator according to this embodiment. [Figure 3]

[0041] This is a cross-sectional side view of a self-generating aircraft compartment thermal indicator of the type shown in Figure 2, according to this embodiment. [Figure 4A]

[0042] This is a cross-sectional side view of a self-generating thermal indicator for an aircraft compartment according to this embodiment. [Figure 4B]

[0043] This is a perspective view of a thermal induction tube of the type shown in Figure 4A, which is incorporated into a self-generating aircraft compartment thermal indicator according to this embodiment. [Figure 4C]

[0044] This is a block diagram showing an aircraft compartment temperature display system for visually detecting, from the outside of the aircraft, a temperature exceeding a threshold in an aircraft compartment located close to the outside of the aircraft, according to this embodiment. [Figure 5A]

[0045] This figure shows a test system and method for a self-generating aircraft compartment thermal indicator according to this embodiment. [Figure 5B]

[0046] This is a diagram of a test system for a self-generating aircraft compartment thermal indicator according to this embodiment. [Figure 6]

[0047] This is a cross-sectional side view of an aircraft component that takes the form of an aircraft wing component equipped with a thermal indicator for a self-generating aircraft compartment according to this embodiment. [Figure 7]

[0048] This is a diagram of the leading edge components of the main wing according to this embodiment. [Figure 8]

[0049] This is a flowchart outlining the method according to this embodiment. [Figure 9]

[0050] This is a flowchart outlining the method according to this embodiment. [Figure 10]

[0051] This is a flowchart outlining the method according to this embodiment. [Modes for carrying out the invention]

[0041]

[0052] For example, it is desirable to detect undesirable levels of heat having elevated temperatures (e.g., which can be demonstrated by quantitative and / or qualitative temperature measurements) above and / or exceeding the expected and / or acceptable temperatures (equivalent herein to “threshold temperatures”), including the detection of elevated temperatures caused by fire and temperatures that may precede the ignition of a fire. According to several embodiments, a self-generating aircraft compartment thermal indicator (equivalent herein to “aircraft compartment thermal indicator,” “aircraft thermal indicator,” and “thermal indicator”) is disclosed. The thermal indicator may be detachably installed in the aircraft's substrate and / or aircraft components (e.g., part of the aircraft's components), including components of the aircraft compartment. According to several embodiments, a self-generating aircraft compartment thermal indicator includes a portion of the thermal indicator that is exposed in the external region of the aircraft and otherwise visible, and visible from an external viewing position, in which case the other portion of the thermal indicator extends into an internal region of the aircraft, which may be a local internal region, and further, at least partially enclosed, an aircraft compartment.

[0042]

[0053] In several of these embodiments, the term “aircraft compartment” includes a fully enclosed area within an aircraft, and further includes an at least partially enclosed interior area within an aircraft. That is, in several of these embodiments, an aircraft compartment may not be a fully and / or fully enclosed compartment. In another embodiment, an aircraft compartment may be a substantially fully enclosed interior compartment of an aircraft.

[0043]

[0054] According to several further embodiments, a “self-generating” thermal indicator means that the thermal indicator comprises an internal and / or self-contained indicator power supply and / or power source that can be activated in the presence of a temperature above a selected threshold temperature. The self-generating aircraft compartment thermal indicator does not realize and / or cause power consumption of the aircraft’s power supply and / or the aircraft’s power “grid” while in operation. That is, the self-generating thermal indicator forms a self-contained battery on demand while in operation, and the thermal indicator operates discretely from the aircraft’s power supply, the aircraft’s auxiliary power supply, etc., without engaging with them and without communicating with them in any other way.

[0044]

[0055] Figure 1 is a representative and non-restrictive diagram of an aircraft of a type that may be equipped with a thermal indicator for the self-generating aircraft compartment, in which the aircraft 10 comprises a fuselage 12 and a main wing 14, the main wing further comprising a leading edge 16, and the main wing 14 housing the leading edge components 16a inside the main wing within a wing compartment housed within the main wing 14.

[0045]

[0056] According to several embodiments, the self-generating aircraft compartment thermal indicator may be configured to be detachably attached to an aircraft component for the purpose of indicating that a threshold temperature selected in at least one of the following has been exceeded: an aircraft compartment located in close proximity to and / or adjacent to the thermal indicator, and a localized internal region of the aircraft located in close proximity to and / or directly adjacent to the temperature-sensing side of the thermal indicator of the aircraft compartment (e.g., a second side of the indicator body), through which the temperature-sensing side of the thermal indicator extends. According to several embodiments, the aircraft compartment may, but does not have to be, a fully enclosed compartment through which the thermal indicator of the aircraft compartment extends. In this embodiment, the aircraft compartment may be a leading-edge compartment of an aircraft wing capable of housing an aircraft bleed duct, which includes an aircraft bleed duct joint located along the bleed duct.

[0046]

[0057] According to several embodiments, when the ambient temperature in an aircraft compartment and / or a localized internal area of ​​the aircraft exceeds a selected threshold temperature, the self-generating thermal indicator of the aircraft compartment may be operablely triggered to form an activated power source (e.g., activated power supply) within the thermal indicator, which is equivalent to a battery as herein. In this case, the activated power supply communicates with a visually detectable indicator (e.g., a light source including flashing light) that can be powered by the activated power supply, and the visually detectable indicator is configured to emit a visually detectable signal that can be detected at an external location of the aircraft outside the aircraft.

[0047]

[0058] According to several embodiments, the self-generating aircraft compartment thermal indicator is configured to respond to a change in ambient temperature from an ambient temperature below a threshold temperature and / or below a threshold temperature range to an elevated temperature above the threshold temperature and / or above a threshold temperature range. That is, the self-generating aircraft compartment thermal indicator may be configured to activate at a selected temperature above a selected threshold temperature. The self-generating aircraft compartment thermal indicator comprises a force-related component, which may be a spring with a potential outward force, when the self-generating thermal indicator is maintained in an environment having an ambient temperature below a threshold temperature and is held in a compressed state by a retainer. Below the threshold temperature, the retainer (as herein equivalently referred to as a disassemblable retainer) has sufficient retaining force on the compressed spring to exceed the potential outward force of the compressed spring, resulting in the spring being held in a compressed or "held" and / or "suppressed" state.

[0048]

[0059] When the thermal indicator of this self-generating aircraft compartment is exposed to and / or exposed to a temperature exceeding a selected threshold temperature, the decomposable retainer, which may be made of a cryogenic alloy, begins to change state and / or continues to experience a significant decrease in retaining force in other ways. Thereafter, the retaining force is overcome, or otherwise exceeded, by the potential outward force of the spring. When this happens, the spring is released from its compressed or held state and expands outward along its longitudinal axis. In this case, the spring can now exert force on an element, for example, an element positioned adjacent to a spring in a compressed state, from the initial stationary position of the element to the position of the element "driven" or activated by the spring.

[0049]

[0060] In this embodiment, during operation, a force ranging from approximately 10 lbf to approximately 30 lbf may be required to rupture a fragile battery electrolyte container enclosed within the thermal indicator. Therefore, the thermal indicator includes a spring configured to apply a similar force ranging from approximately 10 lbf to approximately 30 lbf to the container rupture element (e.g., a piston) when the spring is released. As described herein, to hold the spring element in a compressed state when the thermal indicator of the aircraft compartment is in a “non-activated state,” a disassemblable retainer is configured to be in a solid state and to hold the spring in a compressed or retained state, such that the disassemblable retainer applies a retaining force to the spring that exceeds the outward spring release force.

[0050]

[0061] In this embodiment, the decomposable retainer is made from and / or includes a low-melting-point material (hereinafter equivalent to "low-melting-point material") which may be a low-melting-point metal, a low-melting-point alloy (hereinafter equivalent to "low-melting-point alloy"), a wax, and / or a material that will continue to undergo a phase change from a solid material to a liquid or semi-solid phase at a selected threshold temperature and / or a selected threshold temperature range. That is, according to several embodiments, when the thermal indicator is exposed to an environment having a temperature of at least a selected threshold temperature, the low-melting-point material will continue to undergo a change in physical properties, at least in terms of solid-state retention, and the low-melting-point material will "melt" into a liquid or semi-solid state (e.g., "semi-solid" is a state in which the material begins to flow). In this case, the liquid and semi-solid states are collectively referred to herein as the "liquid state".

[0051]

[0062] In other words, according to several embodiments, the low-melting-point material used to form a decomposable retainer for the thermal indicator of the aircraft compartment is in a solid state, and when the ambient temperature in the aircraft compartment to which the thermal indicator is exposed is below a threshold temperature, which may be, for example, below about 140 degrees Fahrenheit, the decomposable retainer maintains a holding force in a solid state that exceeds the outward potential force of a spring that can impregnate, for example, the low-melting-point material of the decomposable retainer, and the decomposable retainer maintains the spring in a compressed and / or “held” state. In one embodiment, when the ambient temperature to which the thermal indicator is exposed rises to and / or exceeds a threshold temperature in the range of about 140 degrees Fahrenheit to about 200 degrees Fahrenheit, the low-melting-point material used to form the decomposable retainer changes state from a solid phase to a liquid phase (e.g., a liquid “state”). As a result, the low-melting-point material in liquid form (e.g., in a "non-solid" state) then has a coercive force smaller than the outward potential force of the spring (lower than the coercive force when the low-melting-point material is in a solid state), and consequently the spring is released from its initial compressed state into an expanded, released spring state.

[0052]

[0063] As described herein, according to several embodiments, the decomposable retainer includes a “low melting point” material (equivalently referred herein as “low-temperature material” and / or “low-temperature alloy”) in a solid phase below a selected threshold temperature of about 200 degrees Fahrenheit. In another embodiment, the selected threshold temperature is in the range of about 160 degrees Fahrenheit to about 240 degrees Fahrenheit. In another embodiment, the selected threshold is the temperature at which a fire may occur from electronic equipment, in the range of about 160 degrees Fahrenheit to about 180 degrees Fahrenheit. In another embodiment, the selected threshold temperature is the high temperature in an aircraft's hydraulic compartment, in the range of about 160 degrees Fahrenheit to about 280 degrees Fahrenheit. In another embodiment, the selected threshold temperature is in the range of about 140 degrees Fahrenheit to about 165 degrees Fahrenheit. In another embodiment, the selected threshold for cargo storage compartments and / or aircraft cabin storage compartments may be in the range of about 160 degrees Fahrenheit to about 170 degrees Fahrenheit.

[0053]

[0064] In other words, according to several embodiments, the thermal indicator of the aircraft compartment may be selectively tuned or otherwise constructed to provide a temperature indication at a selected temperature and / or selected temperature range, which may vary depending on the location where the temperature indication is installed, and further depending on the area of ​​the aircraft and / or aircraft compartment selected to be monitored for undesirable high temperature changes or presences. According to several embodiments, the temperature indication characteristics of the thermal indicator may be selected to substantially match a threshold temperature monitored on the aircraft, which may include monitoring of aircraft compartments that may require overheat protection (early high temperature detection can be very beneficial). Aircraft compartments that may benefit from overheat protection may include, for example, fuel tanks, areas adjacent to fuel tanks, compartments housing hot air and / or bleed ducts, cargo compartments, passenger storage compartments, compartments housing electrical and / or hydraulic components, battery compartments, engine fan cases, landing gear compartments, and the like.

[0054]

[0065] Further embodiments of this invention consider the detection of threshold temperatures occurring within an aircraft compartment, which are likely to provide evidence of a system anomaly or system failure from expected system function, and which may include, for example, the detection of heat escaping from a duct path (e.g., a junction area of ​​a duct path), including an extraction duct located within the wing compartment at the leading edge of the main wing.

[0055]

[0066] The low-melting-point material may be a metal, a metal alloy, or another low-melting-point material (e.g., wax), and may include one or more of the following: bismuth-containing compounds, lead-containing compounds, tin-containing compounds, indium-containing compounds, thallium-containing compounds, gallium-containing compounds, Rose alloys, Cerosafe, Wood alloys, Field alloys, Cellolow 136, Cellolow 117, gallium, and combinations thereof. The low-melting-point material may further include one or more of the following: wax-containing compounds, other non-metal-containing compounds, and combinations thereof. In this embodiment, the decomposable retainer includes a low-melting-point material having a melting point in the range of about 140°F to about 165°F.

[0056]

[0067] According to this embodiment, at a threshold temperature, when the decomposable retainer undergoes a phase change from solid to liquid, a compressed spring embedded in the decomposable retainer material is released, and the spring is released along its longitudinal axis from a compressed state to an expanded state, and the mechanical force of the spring is applied to a movable element (e.g., a piston) such that it ruptures a fragile container (equivalently referred to herein as the “fragile container,” “fragile battery electrolyte container,” or “fragile battery electrolyte vessel”) containing liquid battery electrolyte. The liquid battery electrolyte is released from the ruptured fragile container, and the electrolyte is driven into the battery compartment by the piston. The presence of the battery electrolyte in the battery compartment thereby forms a battery that can supply current (e.g., charge) to a display element. The display element may be a luminaire, and may be a visually detectable indicator (equivalently referred to as a “visually detectable signaling device”), which may be a light-emitting diode (LED). The current from the activated battery travels along the established electrical circuit from the activated battery to activate the lighting fixture, and a visually detectable ray of light (which may be a ray of light from a flashing lighting fixture) is emitted from the lighting fixture, providing a visual indication of a temperature exceeding a selected threshold temperature that exists within the area of ​​the thermal indicator.

[0057]

[0068] Figures 2 and 3 are cross-sectional side views of a thermal indicator 20 for an aircraft compartment, which is fixedly and detachably attached in place to a base material 18 of an aircraft component, having an outer surface 18a and an inner surface 18b of the base material, and extending through the base material 18. Figures 2 and 3 show the same thermal indicator 20, but the internal components of the thermal indicator 20 are moving from a compressed or compacted state in Figure 2 to an expanded or "released" state in Figure 3.

[0058]

[0069] As shown in Figure 2, the base material 18 of the aircraft component may be within the aircraft component shown as the main wing 14 in Figure 2. As shown in Figure 2, the thermal indicator of this self-generating aircraft compartment comprises an indicator body 22 having an indicator body interior 22a and an indicator body exterior 22b. The indicator body 22 further comprises an indicator body wall 24 bounded by a first side 24a of the indicator body wall (equivalent herein to “indicator body wall interior and / or interior” 24a) and a second side 24b of the indicator body wall (equivalent herein to “indicator body wall exterior and / or exterior” 24b), otherwise extending from them and positioned between them.

[0059]

[0070] The indicator body 22 further comprises a first end 26 of the indicator body which may be substantially flush with the outer surface 18a of the aircraft component. The indicator body 22 further has an indicator body length "l" extending from the first end 26 of the indicator body to a second end 28 of the indicator body. The indicator body 22 further comprises a second end 28 of the indicator body which may extend a distance "d" away from the inner surface 18b of the aircraft component and may extend into the aircraft compartment. As shown in Figure 2, the second end 28 of the indicator body further comprises a second end base 29 of the indicator body. According to several embodiments, the first end 26 of the indicator body is equivalently referred herein to as the "display end" of the indicator, and the second end 28 of the indicator body is equivalently referred herein to as the "temperature sensing end" and / or "sensing end" of the indicator.

[0060]

[0071] In this embodiment (not shown in Figure 2), the first end 26 of the indicator body may extend slightly beyond the outer surface 18a of the aircraft component by a certain distance. In another embodiment (not shown in Figure 2), the first end 26 of the indicator body may countersink into the outer surface 18a of the aircraft component by a selected distance.

[0061]

[0072] As further shown in Figure 2, the first end 26 of the indicator body further comprises a through-opening 26a of the first end of the indicator body, defined by a through-opening wall 26b of the first end of the indicator body. The first end 26 of the indicator body further comprises a first end screw 26c of the indicator body configured to physically engage with a dimensionally matched screw (e.g., a dimensionally "fitting" screw) in the base material 18 of the aircraft component, for the purpose of fixing and releasably (and detachably) engaging the thermal indicator 20 of the aircraft compartment at a selected position in the base material 18 of the aircraft component.

[0062]

[0073] As shown in Figure 2, the self-generating aircraft compartment thermal indicator 20 comprises three sections: 1) a battery compartment 30 bounded by a first end 26 of the indicator body and a battery compartment base 30a, otherwise located adjacent to them; 2) a first chamber 25a of the indicator body defined by the battery compartment 30 and a first side 38a of the piston base 38 of the piston 36, otherwise located adjacent to them and between them; and 3) a second chamber 25b of the indicator body defined by a second side 38b of the piston base 38 of the piston 36 and a second end base 29 of the indicator body at a second end 28 of the indicator body, otherwise located adjacent to them and between them. The battery compartment 30 is shown in Figure 2 comprising a plurality of battery plates 32 and a battery compartment base 30a, the battery compartment base 30a having a through-opening 30b of the battery compartment base.

[0063]

[0074] As further shown in Figure 2, the first chamber 25a of the indicator body is configured to contain, or otherwise house, a fragile container 34 in an unruptured state, the fragile container 34 shown containing a certain amount of battery electrolyte 34a sealed within the fragile container 34 in an unruptured state. Figure 2 further shows an internal piston 36 (equally referred to herein as “piston 36”), the internal piston 36 further comprises a piston base 38, the piston base comprising a first side 38a of the piston base facing, or otherwise forming, the boundary “wall” of the first chamber 25a of the indicator body. The piston base 38 further comprises a second side 38b of the piston base facing, or otherwise forming, the boundary “wall” of the second chamber 25b of the indicator body. As shown in Figure 2, the first chamber of the indicator body may be bounded and / or defined by the battery compartment base 30a, the first side 24a of the indicator body wall, and the first side 38a of the piston base. The piston 36 further comprises a piston rod 39 having a first end 39a of the piston rod that is in direct contact with a luminaire 40, which may be a light-emitting diode (LED) having a light lens 40a. The piston rod 39 further comprises a second end 39b of the piston rod that may be attached to or integrated with the piston base 38 at the first side 38a of the piston base.

[0064]

[0075] As shown in Figure 2, the second chamber 25b of the indicator body comprises a spring 42 having a first end 42a of the spring which can be positioned directly adjacent to the second side 38b of the piston base. The spring 42 further comprises a second end 42b of the spring which can be positioned directly adjacent to the second end base 29 of the indicator body. Figure 2 further shows a disassemblable retainer 44 embedded in the spring 42, or otherwise presented on the spring 42 as a “coating,” “build-up layer,” “impregnation coating,” etc. Thereafter, the disassemblable retainer in a solid state has a form of robustness sufficient to overcome the outward force of the spring 42. That is, according to several embodiments, the spring 42 in a compressed state is held in a compressed state, and in that case, the spring holds its potential outward force while the spring is embedded in the disassemblable retainer 44 while the disassemblable retainer 44 is in a solid state (e.g., “solid”). In the compressed spring state shown in Figure 2, the second chamber 25b of the indicator body may be substantially bounded and / or defined by the material forming the disassemblable retainer 44, the second side 38b of the piston base, and the second end base 29 of the indicator body.

[0065]

[0076] According to several embodiments, as described herein, the decomposable retainer 44 is made from, or otherwise includes, a selected “low melting point” material. The material begins to change from a first state, which is substantially solid (e.g., substantially “solid phase”), having a first retaining force, to a second state. The second state is at least one of a semi-solid state, which may be, for example, a liquid state and a “gelled” state, having a second retaining force smaller than the first retaining force in the liquid state, in which case the second retaining force is also smaller than the potential outward force of a compressed spring that is successfully held in a compressed state (e.g., “held” by a solid retainer) when the decomposable retainer is in the solid phase.

[0066]

[0077] Therefore, Figure 2 shows a “not activated” self-generating aircraft compartment thermal indicator when the ambient temperature is below the threshold temperature (for example, a spring is kept in a compressed state, and insufficient force is applied to the fragile container to rupture it and release its contents). The “not activated” configuration of the thermal indicator means that the thermal indicator's battery is not formed in the thermal indicator.

[0067]

[0078] In contrast to Figure 2, according to multiple embodiments, Figure 3 shows a self-powered aircraft compartment thermal indicator 20 now in an activated configuration in the presence of a high temperature above a selected threshold temperature (e.g., a temperature greater / higher than a lower initial ambient temperature). The “activated” configuration of the thermal indicator means that the thermal indicator’s battery is formed within the thermal indicator and supplies power to the thermal indicator.

[0068]

[0079] As shown in Figure 3, in detecting a temperature above a selected threshold temperature (represented by the presence of a heat source 46 that can take the form of a flame and cause high temperatures), the thermal indicator 20 is "activated" to an activated indicator state (from the unactivated indicator state shown in Figure 2), in which case the disassemblable retainer 44 now moves from the solid phase to the liquid phase, releasing the spring 42 from the compressed spring state (shown in Figure 2). In the activated indicator state shown in Figure 3, the potential outward force of the spring 42 exerts an realized unsuppressed and / or unrestricted outward actual force on the piston 36 along the longitudinal axis of the spring (shown as an "upward" force in Figure 3). As shown in Figure 3, the area of ​​the second chamber 25b of the indicator body expands to substantially occupy, or otherwise increase, the area of ​​the first chamber 25a of the previous indicator body that was present in the thermal indicator when the thermal indicator was in an "unactivated" state (the first chamber 25a of the indicator body is shown in Figure 2, but not in Figure 3).

[0069]

[0080] During operation, when the thermal indicator 20 is exposed to a temperature above the selected threshold temperature, a force from the release spring 42 is transmitted to the piston 36, pushing the piston 36 into the first chamber 25a of the indicator body, in which case sufficient force pushes the piston to compress the fragile container 44 between the piston base 38 and the battery compartment base 30a, and sufficient force ruptures the fragile container 34. The rupture of the fragile container 34 releases the battery electrolyte 34a from the now ruptured fragile container 34 (which may be an ampoule, for example, a thin-walled glass ampoule). Once the liquid electrolyte 34a is released from the now ruptured fragile container 34, the electrolyte is driven from the shrinking area of ​​the first chamber of the indicator body through the opening 30b in the battery compartment base 30a and into the battery compartment 30.

[0070]

[0081] As shown in Figures 2 and 3, the O-ring 38c is positioned around the piston base 38 to form a seal that is substantially leak-proof between the piston base 38 and the first side of the indicator body (e.g., the "inner surface" of the indicator body wall). On the other hand, the O-ring 38c also facilitates, and otherwise does not hinder, the relative movement of the piston in response to the force of the spring 42 (e.g., a force moderate enough to burst the fragile container and effectively release the battery electrolyte from the fragile container into the battery). In addition to preventing unwanted electrolyte leakage into the second chamber of the indicator body, the O-ring 38c is made from a resilient and robust material (e.g., silicone, rubber, etc.) and has a selected Young's modulus and other O-ring material properties (e.g., physical, chemical properties, etc.) that help to keep the electrolyte in place within the battery compartment 30.

[0071]

[0082] As shown in Figure 3, the position of the piston base 38 in direct contact with the battery compartment 30a further seals the through-opening 30b of the battery compartment base, in which case the maximum available amount of electrolyte 34a (hereinafter equivalently referred to as “battery electrolyte”) is transferred from the ruptured fragile container 34 (containing the electrolyte) into the battery compartment 30, reacts with the battery plate 32 to form a now activated and / or “activated” working battery 33, which is configured to generate current along the battery lead 33a, activating the luminaire 40, which emits a ray of light 41 from the luminaire 40 (which may be, for example, an LED) in an operable (and / or “activated” and / or “activated”) thermal indicator state.

[0072]

[0083] As further shown in Figure 3, similar to the sealing function of the piston base O-ring 38c, the piston rod O-ring 38d is positioned in the through-opening wall 26b of the first end of the indicator body and configured to contact the emerging piston rod 39. The piston rod O-ring 38d is further configured and positioned to form a substantially leak-free seal to facilitate the retention of the battery electrolyte within the battery 33 and within the battery compartment 30. Similar to the piston base O-ring 38c, the piston rod O-ring 38d is made from and / or may contain resilient and robust materials (e.g., silicone, rubber, etc.). These materials may have a selected Young's modulus and other O-ring material properties (e.g., physical, chemical properties, etc.). This helps to maintain the electrolyte in place within the battery compartment 30.

[0073]

[0084] Figures 4A and 4B show further embodiments of the present invention, including a thermal indicator 20 of the self-generating aircraft compartment. The indicator 20 may further comprise one or more integrated heat transfer elements. These heat transfer elements, according to this embodiment, further enhance the effectiveness and sensitivity of the thermal indicator by concentrating and conserving the effects and influences of temperature increases occurring in close proximity to the thermal indicator.

[0074]

[0085] Figure 4A is a cross-sectional side view of a thermal indicator 20 for an aircraft compartment, fixedly and detachably attached to a base material 18 of an aircraft component, having an outer surface 18a of the aircraft component's base material and an inner surface 18b of the aircraft component's base material, and extending through the base material 18. Figure 4A shows a self-generating aircraft compartment thermal indicator 20 of the type shown in Figure 2, further comprising similarly numbered components. Figure 4A shows a thermal indicator 20 further comprising an integrated heat transfer element taking the form of a heat induction tube 50. The heat induction tube 50 is located within a second chamber 25b of the indicator body of the thermal indicator 20, and otherwise integrated (and can be fixedly attached). In that case, the heat induction tube 50 extends into the second chamber 25b of the indicator body and comprises a first end 50a of the heat induction tube that is positioned such that the outer circumference of the heat induction tube wall 54 is located close to and / or substantially directly adjacent to a disassemblable retainer 44. In this embodiment, the heat induction tube 50 is in direct contact with the decomposable retainer 44. The second end 50b of the heat induction tube is shown in Figure 4A. The second end 50b extends beyond the second end base 29 of the indicator body by a selected distance from the second chamber 25b of the indicator body. As shown in Figure 4A, the heat induction tube 50 may further comprise a “hot air” vent 66 located at the first end 50a of the heat induction tube. The “hot air” vent 66 can further guide, circulate, or otherwise distribute heat to the decomposable retainer 44.

[0075]

[0086] Figure 4B shows a perspective view of a segment of a thermal induction tube 50 of the type shown in Figure 4A. The thermal induction tube 50 may be incorporated into the thermal indicator 20 shown in Figure 4A. As also shown in Figure 4B, the thermal induction tube 50 may be substantially cylindrical and have an outer circumference sized such that the thermal induction tube substantially "fills" or otherwise substantially occupies the available area present in the second chamber of the indicator body. In this case, the outer surface of the edge of the tube fin 52 extends outward from the outer surface of the thermal induction tube wall 54 of the thermal induction tube 50 (forming the outer circumference of the thermal induction tube 50). The outer surface of the edge of the tube fin 52 may be in direct contact with a degradable retainer 44 containing a cryogenic material used to manufacture the degradable retainer element. According to several embodiments, the outer edge of the tube fin 52 is referred herein equivalently to the outer circumference of the thermal induction tube wall. Furthermore, several embodiments also consider the incorporation of a thermal induction tube without the tube fin 52. In that case, the thermal induction tube wall is then positioned close to and adjacent to the disassemblable retainer. In another embodiment, the outer circumference of the thermal induction tube wall is positioned in direct contact with the disassemblable retainer (for example, without the tube fins 52).

[0076]

[0087] As shown in Figures 4A and 4B, the heat induction tube 50 can be “cut out” to further include a heat induction tube central path 56 extending along the length of the heat induction tube 50, in which case the heat induction tube central path 56 is defined by the inner surface of the heat induction tube inner wall 54.

[0077]

[0088] As further shown in Figures 4A and 4B, the tube fins 52 extend outward from the inner wall 54 of the thermal induction tube by a selected distance, in which case the tube fins 52 are configured to facilitate the radiation of heat that can be absorbed by the thermal induction tube 50 (e.g., absorbed by the thermal induction tube of a thermal indicator in an aircraft compartment), which is then directed to and / or "radiated" to a decomposable retainer 44 located in the second chamber 25b of the indicator body of the thermal indicator 20. As shown in Figure 4A, the second end 50b of the thermal induction tube extends from the second end base 29 of the thermal indicator body to enhance the collection of heat that may be present and generated outside and in the vicinity of the thermal indicator 20 and outside and in the vicinity of the second end 50b of the thermal induction tube (heat that can be accumulated in an aircraft compartment to which the second end of the indicator body may extend, and heat that exposes the second end of the thermal indicator to the aircraft compartment environment and the temperature present in the aircraft compartment environment). These heat induction tubes may be made from materials having high heat transfer capacity, including, for example, copper, copper alloys, and combinations thereof, as well as metals and metal alloys.

[0078]

[0089] According to several embodiments, the power source for operating the activated self-generating aircraft compartment thermal indicator resides solely within the thermal indicator and battery, which is formed on demand within the thermal indicator by combining battery electrolyte released from a ruptured, fragile container with a battery plate within the thermal indicator's battery compartment, in the presence of a threshold temperature. According to several embodiments, the battery type formed within the thermal indicator may be a fuse-type battery.

[0079]

[0090] When activated and operational, the self-generating aircraft compartment thermal indicator is configured to supply a visually detectable signal to personnel at or very close to the temperature detector located at a position directly outside the monitored aircraft compartment, depending on the intensity of one or more rays of light emitted from the illuminator. In other words, the thermal indicator presents a significant advantage for ground personnel monitoring a potentially overheating aircraft compartment. In this case, it provides a visual indication of an overheating compartment visible to ground personnel from outside the aircraft when the aircraft is on the ground, in a non-flight phase, and personnel may not be located in the cockpit or may not have access to the cockpit flight control panel. According to this embodiment, the flashing light emitted from the thermal indicator's illuminator, powered by the thermal indicator of the self-generating aircraft compartment, can visually indicate and warn personnel of the presence of excessive temperature (e.g., temperature exceeding a selected threshold), and can visually communicate to ground personnel that an internal compartment located within the aircraft, or an internal compartment located in an area immediately adjacent to a flashing thermal indicator, is experiencing, or otherwise maintaining, an excessive temperature requiring attention (e.g., by viewing a visual display or being warned via a visual "temperature alarm").

[0080]

[0091] Therefore, this self-generating aircraft compartment thermal indicator, configured to transmit temperature information at the location of potential thermal events, is in stark contrast to typical aircraft temperature sensing systems. Typical aircraft temperature sensing systems may be connected in series, connected to the aircraft's power supply and other sensors via extensive wiring, or otherwise communicating, and further require associated hardware, software, mechanical coupling and mounting, processors, readers, control panels, and other centralized system components. This may relay the detected temperature information to a centrally located location within the aircraft (e.g., the cockpit panel), adding significant costs and weight to the aircraft, further increasing the aircraft's operating costs, otherwise negatively impacting the aircraft's flyable range, increasing the aircraft's fuel consumption, and reducing the aircraft's available passenger seating capacity.

[0081]

[0092] According to several embodiments, once activated, the activated self-generating aircraft compartment thermal indicator cannot return to an unactivated state once a disassemblable retainer is disassembled, the spring is released, and force is applied to the piston, causing the fragile container holding the battery electrolyte to rupture. According to further embodiments, the thermal indicator is detachably fixed and positioned within the aircraft structure. Thereafter, the activated thermal indicator can be removed from the aircraft structure and replaced with another "unactivated" thermal indicator.

[0082]

[0093] Figure 4C is a box diagram outlining a system for displaying temperatures exceeding a threshold temperature in an internal region of an aircraft, positioned in close proximity to an external location of the aircraft, and for identifying the presence of such temperatures. According to several embodiments, as shown in Figure 4C, the system 58 comprises an aircraft 10. The aircraft 10 further comprises an aircraft compartment substrate 18. The substrate 18 further comprises a self-generating aircraft compartment thermal indicator 20 of the type described herein.

[0083]

[0094] Figures 5A and 5B illustrate several present embodiment test methods (e.g., inspection and / or quality control protocol methods) for testing the force loading / release / activation of the thermal indicator (Figure 5A), and for testing the satisfactory operation and / or functional operation of the luminaire element of the thermal indicator in situ, in its installed state, and / or during storage (Figure 5B), including when the thermal indicator's battery is not activated.

[0084]

[0095] As shown in Figure 5A, a system 60 (equivalently referred to herein as the load cell test system 60) adopting the present test method for the thermal indicator is shown to determine the spring release force that a decomposable retainer will release at a threshold temperature, and to test that when the thermal indicator is in the presence of or otherwise exposed to a high temperature representing a temperature above a selected threshold temperature, the spring will release as intended and rupture a fragile battery electrolyte container with an appropriate applied force supplied to the piston. As shown in Figure 5A, a load cell 64 (equivalently referred to herein as the force gauge 64) may be placed in a cap 62 that fits over the thermal indicator 20. In this case, the load cell is configured to evaluate, or otherwise measure, the releaseable or potential force of the spring 42. The load cell 64 is in communication with a data acquisition device / processor 65 via a data lead 64a and includes a display 68 configured to relay signals and information generated by the load cell and interpreted by the processor via a data line 67 (e.g., to an observer, technician, etc.).

[0085]

[0096] According to this test protocol for testing the operating force of the thermal indicator 20 of this self-generating aircraft compartment, a test fixture is supplied to support the thermal indicator, and a data acquisition device 65 equipped with a processor / computer is mounted on the test fixture. Data acquisition software is run to record and display the data acquired through the test. A temperature sensor is positioned near the temperature element of the thermal induction element, and it can be ensured that the thermal indicator is properly calibrated and connected to the data acquisition device. Temperature changes are established to verify the operation of the temperature sensor function by heating or cooling the sensor. A force gauge is positioned to contact the first end of the thermal indicator and further brought to contact the first end of the piston rod in the illuminator (when not activated, the force gauge is configured to remain substantially flush with the first end of the thermal indicator body). The force gauge is set to "zero" or otherwise calibrated according to the manufacturer's instructions. For example, a heat source, which may take the form of a heat gun 46 (equivalently referred to as a "hot air gun"), is positioned to provide heat to the thermal indicator and the thermal indicator force test / verification system. The heat is applied to the thermal induction tube of the thermal indicator, in which case the heat induction from the heat source is uniformly distributed across the thermal induction tube of the thermal indicator. The heating time is marked and recorded using a timing device (e.g., a stopwatch). When the threshold temperature of the thermal indicator is approached and reached, the time is marked, and the internal spring releases the generated force. The generated force is recorded in a force gauge, in which case the maximum force applied to the force gauge during spring release may be displayed on the system display, or otherwise displayed and recorded. The time of temperature and spring release force may be recorded and plotted.

[0086]

[0097] Figure 5B shows an in situ test of a self-generating aircraft compartment thermal indicator 20 of the present disclosure, which is installed in a fixed and detachable position within a substrate 18 of an aircraft component. According to the test system 70, for example, to evaluate the proper operating condition of the thermal indicator during scheduled maintenance and / or inspection of an aircraft component equipped with the thermal indicator, the positive ("+") terminal of the battery 72 is placed in contact with an exposed pad of the thermal indicator via a test lead 76, and the negative ("-") terminal of the battery 72 is placed in contact via a test lead 74. When the current supplied by the battery 72 powers the thermal indicator, flashing light in the illuminator 40 indicates that the illuminator is functioning properly, confirming that the circuit in the thermal indicator is working.

[0087]

[0098] Figure 6 is a typical cross-sectional side view of a section of the leading edge 16 of an aircraft wing 14. The leading edge 16 may be of the type of aircraft component substrate 18 described herein and may include the outer surface 18a of the aircraft component substrate. The wing leading edge 16 as shown in Figure 6 further comprises a wing leading edge compartment 16a. The compartment 16a may further comprise a wing leading edge bleed duct 16b (shown in Figure 7) having a plurality of wing leading edge bleed duct joints 16c inside the compartment 16a. Figure 6 further shows a self-generating aircraft compartment thermal indicator 20 detachably attached to the aircraft component substrate, in which case the first end of the indicator body is at least partially visible on the outer surface 18a of the aircraft component substrate 18 shown as the wing leading edge component 16. In other words, as shown in Figure 6, the thermal indicator 20 is positioned such that the first end 26 of the “display” indicator body is visible on the exterior of the outer surface 18a of the substrate 18 of the aircraft component, and the second end 28 of the “temperature sensing” indicator body is positioned within, or otherwise extending into, the compartment 16a of the leading edge of the aircraft's main wing.

[0088]

[0099] Figure 7 is a perspective view of a leading-edge component 16d of the wing of an aircraft of the type shown in Figure 6, configured to support a leading-edge bleed duct 16b of the wing within a leading-edge compartment 16a of the wing. The leading-edge bleed duct 16b of the wing may be configured to receive and guide hot bleed air along a path that safely disperses the hot bleed air to an outlet (not shown). In the event that one or more of the multiple hot bleed duct joints 16c wear out through use, or otherwise hot bleed air leaks out of the duct (e.g., hot bleed air traversing the leading-edge bleed duct 16b of the wing), the hot bleed air escapes from the duct, enters and occupies the leading-edge compartment 16a of the wing, and significantly increases the temperature of the ambient air within the leading-edge compartment 16a of the wing leading edge 16.

[0089]

[0100] In some embodiments, as shown in Figure 6, in the event that a warm and / or hot bleed air flow 17 leaks from the bleed air duct 16b at the leading edge of the main wing, the second end 28 of the indicator body is exposed to the increased temperature in the compartment 16a at the leading edge of the main wing. When the temperature of the air in the compartment at the leading edge of the main wing rises to a selected acceptable threshold temperature (e.g., in the range of approximately 140 to 165 degrees Fahrenheit), the self-generating aircraft compartment thermal indicator 20 is activated when a disassemblable retainer releases an internal spring, driving an internal piston to a certain position. Once sufficient spring force is released, this causes the battery electrolyte container to rupture, releasing the battery electrolyte from the container and forming an activated battery within the indicator, which powers a luminaire in the first end of the indicator, which is visible or otherwise visually detectable outside the aircraft compartment, indicated as the leading edge 16 of the main wing (e.g., visually detectable by ground personnel outside the aircraft and in an "outside" position). In several embodiments, during operation, in the event of such an alarm or “visual warning” (e.g., flashing light) event emitted from the activated thermal indicator 20, any person who visually detects the flashing light of the thermal indicator 20 will become aware of an undesirable temperature rise within the compartment of the aircraft component housing the thermal indicator (e.g., in this case, the compartment 16a at the leading edge of the main wing).

[0090]

[0101] According to several embodiments, a self-generating aircraft compartment thermal indicator can be detachably engaged within the aircraft's components. In this case, the visual indication of temperature events allows for maintenance, inspection, rework, and replacement after the temperature event to be performed, for example, by ground personnel on the aircraft's exterior. In addition, the removal of the thermal indicator provides access for inspection into the aircraft compartment housing the thermal indicator, resulting in, for example, more efficient and cost-effective inspection of the aircraft compartment by ground personnel from the aircraft's exterior.

[0091]

[0102] In addition, the “single-use” nature of this thermal indicator, which allows for the easy replacement of detachably mounted, self-generating aircraft compartment thermal indicators, further considers the favorable organization, or otherwise significant simplification, of maintaining a stock of various thermal indicators for various locations on the aircraft, which can be installed on various aircraft components on the aircraft. In this case, various types of thermal indicators having different “trigger” or threshold temperatures among various such thermal indicators can be targeted, or otherwise adapted for specified uses and locations on the aircraft. That is, each “type” of this thermal indicator having different trigger threshold temperatures can be distinguished by their intended use in their respective locations, or otherwise classified to be activated at different threshold temperatures depending on the destination of the final location on the aircraft. For example, according to several embodiments, thermal indicators having different activation threshold temperatures can be individually identified through easily identifiable markings, which may include the following: In other words, for example, "color coding" or other readily identifiable features (to ensure that a particular thermal indicator is used correctly in a particular aircraft location according to a threshold temperature established for a compartment of a particular aircraft component) can facilitate and / or organize the maintenance of a stock of available and easily identifiable thermal indicators.

[0092]

[0103] Figures 8, 9, and 10 are flowcharts outlining the method. Figure 8 outlines the method 100 for detecting a threshold temperature outside the aircraft in an internal region of an aircraft component, which may include an aircraft compartment located in close proximity to the outside of the aircraft. In this case, the method 100 includes placing a self-generating aircraft compartment thermal indicator 102 inside the aircraft component. In this case, the aircraft component includes the outer surface of the aircraft component (e.g., the outer surface of the aircraft). In this case, the outer surface of the aircraft component is located adjacent to an internal region of the aircraft component, which may include an aircraft compartment. In this case, the aircraft compartment thermal indicator 20 comprises an indicator body 22 having an indicator body length ("l"). The indicator body further comprises an indicator body interior 22a, an indicator body exterior 22b, a first end 26 of the indicator body (e.g., the display side), a second end 28 of the indicator body (e.g., the temperature sensing side), a second end base 29 of the indicator body, and an internal piston 36. The internal piston is a piston base 38, further comprising a piston base having a first side 38a and a second side 38b of the piston base, and the internal piston further comprises a piston rod 39, having a first end 39a and a second end 39b of the piston rod, the second end of the piston rod being at least connected to the first side 38a of the piston base. The indicator body is an indicator body wall 24, further comprising an indicator body wall, extending along the length of the indicator body from a first end 26 to a second end 28 of the indicator body, with a second side (e.g., indicator body wall "outer surface") 24b on the outside 22b of the indicator body and a first side (e.g., indicator body wall "inner surface") 24a on the inside 22a of the indicator body.The indicator body further comprises a battery compartment 30 comprising a plurality of battery plates and a battery compartment base 30a, and a battery compartment defined by the battery compartment base 30a and the inner surface 24a of the indicator body wall. The indicator body further comprises a first chamber 25a of the indicator body defined by the inner surface 24a of the indicator body wall, a first side 38a of the piston base, and the battery compartment base 30a, and comprising a fragile container 34 for holding a certain amount of battery electrolyte 34b. According to several embodiments, the fragile container is equivalently referred to herein as “fragile container”.

[0093]

[0104] The indicator body further comprises a second chamber 25b of the indicator body, defined by the inner surface 24a of the indicator body wall and the second side 38b of the piston base, and further comprising a spring 42, the spring having a first end 42a fixedly attached to the piston base 38, and a second end 42b positioned directly adjacent to the base 29 of the indicator body. The second chamber of the indicator body further comprises a disassemblable retainer 44, which is in contact with the spring 42 and configured to hold the spring 42 in a compressed spring state within a first temperature range, and to disassemble at a threshold temperature within a second temperature range, and further configured to hold the spring in a compressed spring state while within the first temperature range, and further configured to release the spring from a compressed spring state to an expanded spring state at a threshold temperature within the second temperature range, the threshold temperature being higher than the first temperature range.

[0094]

[0105] Method 100 further includes: causing a decomposable retainer to decompose in the presence of a compartment temperature exceeding a threshold temperature 104; releasing a spring from a compressed spring state to an extended spring state 106; advancing a piston from an initial piston position to a piston deployed position 108; rupturing a fragile container to form a ruptured fragile container 110; releasing at least a portion of a certain amount of battery electrolyte from the ruptured fragile container 112; introducing at least a portion of a certain amount of battery electrolyte into a battery compartment to form a battery 114, the battery being configured to generate and supply current; leading a first end of a piston rod through a through-opening in the first end of an indicator body to a selected distance beyond the outer surface of an aircraft component 116; and introducing current from the battery to a light-emitting device to emit light from the light-emitting device 118.

[0095]

[0106] Figure 9 outlines Method 200, which is outlined in Method 100, shown in Figure 8, and has the features disclosed herein. In this case, Method 200 further includes visually detecting light emitted from a light-emitting device outside the aircraft.202 In this case, the visual detection is located adjacent to or otherwise directly connected to a thermal indicator in the aircraft compartment to demonstrate the presence of at least a threshold temperature occurring within the aircraft compartment.

[0096]

[0107] Figure 10 outlines Method 300, which is outlined in Methods 100 and 200, and shown in Figures 8 and 9, respectively, and has the features described herein. In this case, Method 300 further includes detachably locating a self-generating aircraft compartment thermal indicator within an aircraft component 302.

[0097]

[0108] The method outlined in Figures 8, 9, and 10 and described herein can implement the self-generating aircraft compartment thermal indicator described herein and shown in at least Figures 2, 3, 4A, 4B, 5A, 5B, and 6.

[0098]

[0109] As used herein, the term “substantially” means that certain physical elements, physical arrangements, physical shapes, orientations, etc., are realized almost completely or substantially. That is, for example, according to some of these embodiments, a “substantially” completely enclosed compartment includes a compartment that is substantially and / or almost entirely and completely enclosed.

[0099]

[0110] Several embodiments of this disclosure can, of course, be implemented in ways different from those specifically described herein, as long as they do not deviate from the essential characteristics of this disclosure. These embodiments should be considered in all respects as illustrative and non-limiting, and all modifications that fall within the meaning and equivalence of the claims are intended to be encompassed within the claims.

Claims

1. A self-generating thermal indicator (20) for an aircraft compartment, The indicator body (22) having an indicator body length is provided, and the indicator body is Inside the indicator body (22a), External indicator body (22b), The first end (26) of the indicator body, The second end (28) of the indicator body, Second end base (29) of the indicator body, An internal piston (36) comprising a piston base (38) having a first side (38a) and a second side (38b) of the piston base, and a piston rod (39) having a first end (39a) and a second end (39b) of the piston rod, wherein the first end of the piston rod is connected to a visually detectable signal device (40) and the second end of the piston rod is connected to at least the first side (38a) of the piston base, An indicator body wall (24) that extends along the length of the indicator body from a first end (26) of the indicator body to a second end (28) of the indicator body, and comprises an outer surface (24b) of the indicator body wall on the outside (22b) of the indicator body and an inner surface (24a) of the indicator body wall on the inside (22a) of the indicator body, A battery compartment (30) comprising a battery compartment base (30a), the battery compartment defined by the battery compartment base (30a) and the inner surface of the indicator body wall (24a), A first chamber (25a) of the indicator body, defined by the inner surface (24a) of the indicator body wall, the first side (38a) of the piston base, and the battery compartment base (30a), comprising a fragile container (34) for containing a certain amount of battery electrolyte (34a), A second chamber (25b) of the indicator body defined by the inner surface (24a) of the indicator body wall and the second side (38b) of the piston base, comprising a spring (42) having a first end (42a) of the spring fixedly attached to the piston base (38), and further comprising a second end (42b) of the spring positioned directly adjacent to the second end base (29) of the indicator body, and A disassemblable retainer (44) is connected to the spring (42) and configured to hold the spring (42) in a compressed spring state within a first temperature range, configured to disassemble at a threshold temperature within a second temperature range, further configured to hold the spring in a compressed spring state while within the first temperature range, and further configured to release the spring from the compressed spring state to an expanded spring state at the threshold temperature within the second temperature range. The threshold temperature is higher than the first temperature range, and is a self-generating thermal indicator for an aircraft compartment.

2. The self-generating thermal indicator for an aircraft compartment according to claim 1, wherein the visually detectable signaling device (40) is at least one of a lighting fixture and a light-emitting diode.

3. The thermal indicator for a self-generating aircraft compartment according to claim 1, wherein the battery compartment further comprises a plurality of battery plates (32).

4. The self-generating thermal indicator for an aircraft compartment according to claim 1, wherein the decomposable retainer (44) comprises at least one of a low-temperature alloy and a wax, and the at least one of the low-temperature alloy and the wax and the decomposable retainer have a melting point in the range of about 140 degrees Fahrenheit to about 165 degrees Fahrenheit.

5. The low-temperature alloy comprises a low-temperature metal alloy containing at least one of bismuth, lead, tin, indium, cadmium, thallium, gallium, and combinations thereof, wherein the low-temperature alloy has a melting point in the range of about 140 degrees Fahrenheit to about 165 degrees Fahrenheit, as described in claim 4, for a self-generating aircraft compartment thermal indicator.

6. The self-generating thermal indicator for an aircraft compartment according to claim 5, wherein the low-temperature metal alloy includes at least one of Rose alloy, Cerosafe, Wood alloy, Fields alloy, Cellolow 136, Cellolow 117, gallium, and combinations thereof.

7. The self-generating thermal indicator for an aircraft compartment according to claim 3, wherein the plurality of battery plates (32) in the presence of the electrolyte (34a) are configured to form a battery (33) configured to supply current.

8. The self-generating thermal indicator for an aircraft compartment according to claim 7, wherein the battery is a fuse battery.

9. A self-generating thermal indicator for an aircraft compartment according to claim 1, further comprising a thermal induction tube (50) comprising a first end (50a) of the thermal induction tube, a second end (50b) of the thermal induction tube, and a thermal induction tube wall (54), wherein the first end of the thermal induction tube is integrated into a second chamber of the indicator body, and the thermal induction tube wall is positioned directly adjacent to the disassemblable retainer.

10. Component of an aircraft wing (14)(16)(16a)(16b)(16c) comprising a self-generating aircraft compartment thermal indicator as described in claim 1.

11. An aircraft (10) comprising a self-generating aircraft compartment thermal indicator as described in claim 1.

12. An aircraft compartment temperature display system (58) for visually detecting, from the outside of the aircraft, a temperature in an aircraft compartment close to the outside of the aircraft that exceeds a threshold, The aircraft comprises (14) (16) (16a) (16b) (16c) and a detachably attached self-generating aircraft compartment thermal indicator (20), wherein the self-generating aircraft compartment thermal indicator is The indicator body (22) having an indicator body length ("l") is provided, and the indicator body is Inside the indicator body (22a), External indicator body (22b), A first end (26) of an indicator body visible from the outside of the aircraft, A second end (28) of the indicator body is located within the aircraft compartment and extends into the aircraft compartment by a distance selected from the first end of the thermal indicator. Second end base (29) of the indicator body, An internal piston (36) comprising a piston base (38) having a first side (38a) and a second side (38b) of the piston base, and a piston rod (39) having a first end (39a) and a second end (39b) of the piston rod, the second end of the piston rod being at least connected to the first side (38a) of the piston base, An indicator body wall (24) that extends along the length of the indicator body from a first end (26) of the indicator body to a second end (28) of the indicator body, and comprises an outer surface (24b) of the indicator body wall on the outside (22b) of the indicator body and an inner surface (24a) of the indicator body wall on the inside (22a) of the indicator body, A battery compartment (30) configured to house a battery (33), wherein the battery is configured to supply current along an electrical circuit, the battery compartment comprises a plurality of battery plates (32), a battery compartment base (30a), and the battery compartment is defined by the battery compartment base (30a) and the inner surface of the indicator body wall (24a), A first chamber (25a) of the indicator body, defined by the inner surface (24a) of the indicator body wall, the first side (38a) of the piston base, and the battery compartment base (30a), comprising a fragile container (34) for containing a certain amount of battery electrolyte (34a), A second chamber (25b) of the indicator body, defined by the inner surface (24a) of the indicator body wall and the second side (38b) of the piston base, comprising a spring (42) having a first end (42a) of the spring fixedly attached to the piston base (38), and further comprising a second end (42b) of the spring positioned directly adjacent to the second end base (29) of the indicator body, A disassemblable retainer (44) that is connected to the spring (42), and is configured to hold the spring (42) in a compressed spring state within a first temperature range, and to disassemble at a threshold temperature within a second temperature range, and is further configured to hold the spring in a compressed spring state while within the first temperature range, and is further configured to release the spring from the compressed spring state to an expanded spring state at the threshold temperature within the second temperature range, and A visually detectable signaling device (40) located at the first end of the indicator body, comprising a visually detectable signaling device that communicates with the electrical circuit and is further configured to be visually detectable outside the aircraft, The signal device (40) is configured to activate when the temperature exceeds a threshold temperature, in an aircraft compartment temperature display system.

13. The aircraft compartment temperature display system according to claim 12, wherein the signaling device includes at least one of a lighting fixture and a light-emitting diode.

14. The aircraft compartment temperature indicator system according to claim 12, wherein the signaling device is configured to emit a ray of light in response to the second end of the indicator being exposed to a threshold temperature in the range of approximately 140 degrees Fahrenheit to approximately 165 degrees Fahrenheit.

15. The aircraft compartment temperature display system according to claim 12, wherein the thermal indicator of the aircraft compartment is a detachable aircraft compartment thermal indicator.

16. The aircraft compartment temperature display system according to claim 12, wherein the battery is a fuse battery, and the fuse battery is configured to activate at the threshold temperature.

17. A method (100) for displaying a temperature exceeding a threshold temperature in a compartment of an aircraft component from outside the aircraft, (102) The aircraft compartment thermal indicator is located within an aircraft component, the aircraft component includes the outer surface of the aircraft component, the outer surface of the aircraft component is located adjacent to the aircraft compartment, and the aircraft compartment thermal indicator is located within an aircraft component. The indicator body (22) having an indicator body length (l) is provided, Inside the indicator body (22a), External indicator body (22b), The first end (26) of the indicator body (display side), The through-opening (26a) at the first end of the indicator body, The second end (28) of the indicator body, Second end base (29) of the indicator body, An internal piston (36) comprising a piston base (38) having a first side (38a) and a second side (38b) of the piston base, and a piston rod (39) having a first end (39a) and a second end (39b) of the piston rod, wherein the second end of the piston rod is at least connected to the first side (38a) of the piston base, and the first end of the piston rod is equipped with a light-emitting device (40), An indicator body wall (24) extending along the length of the indicator body from a first end (26) of the indicator body to a second end (28) of the indicator body, to a second end base (29) of the indicator body, and comprising an indicator body wall outer surface (24b) on the outside (22b) of the indicator body and an indicator body wall inner surface (24a) on the inside (22a) of the indicator body, A battery compartment (30) comprising a plurality of battery plates, the battery compartment comprising a battery compartment base (30a), and the battery compartment defined by the battery compartment base (30a) and the inner surface of the indicator body wall (24a), A first chamber (25a) of the indicator body, defined by the inner surface (24a) of the indicator body wall, the first side (38a) of the piston base, and the battery compartment base (30a), comprising a fragile container (34) for containing a certain amount of battery electrolyte (34a), A second chamber (25b) of the indicator body defined by the inner surface (24a) of the indicator body wall and the second side (38b) of the piston base, comprising a spring (42) having a first end (42a) of the spring fixedly attached to the piston base (38), and further comprising a second end (42b) of the spring positioned directly adjacent to the second end base (29) of the indicator body, and A disassemblable retainer (44) is provided, which is connected to the spring (42), and is configured to hold the spring (42) in a compressed spring state within a first temperature range, and is configured to disassemble at a threshold temperature within a second temperature range, and is further configured to release the spring from the compressed spring state while within the first temperature range to an expanded spring state at the threshold temperature within the second temperature range, wherein the threshold temperature is higher than the first temperature range. The aforementioned method further, Decomposing the decomposable retainer in the presence of a compartment temperature exceeding the threshold temperature (104), Releasing the spring from the compressed spring state to the expanded spring state (106), (108) Sending the piston from the initial piston position to the extended piston position, In order to form a ruptured, fragile container, the fragile container is ruptured (110), (112) Releasing at least a portion of the amount of battery electrolyte from the ruptured, fragile container. To form the battery, introduce at least a portion of the aforementioned amount of battery electrolyte into the battery compartment (114), wherein the battery is configured to generate and supply current, introduce at least a portion of the aforementioned amount of battery electrolyte into the battery compartment Leading the first end of the piston rod through the through-opening of the first end of the indicator body to a selected distance beyond the outer surface of the aircraft component (116), and A method (100) comprising (118) drawing current from the battery to activate the light-emitting device in order to emit light from the light-emitting device.

18. The method according to claim 17 (200), further comprising visually detecting the light emitted from the light-emitting device outside the aircraft (202).

19. The method according to claim 18 (300), further comprising (302) detachably arranging the thermal indicator of the aircraft compartment within the aircraft component.

20. The method according to claim 18 (200), wherein visually detecting illumination light from a thermal indicator in the aircraft compartment outside the aircraft is proof that the temperature inside the aircraft's components has exceeded at least a threshold temperature.

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