Abstract:
The present invention concerns a device for use in an aircraft fuel pressure sensor line. More particularly, but not exclusively, this invention concerns a fuel pressure sensor line connecting an aircraft fuel pump to a pressure sensor. The invention also concerns a fuel pressure sensor line reservoir for use with a sensor line connecting an aircraft fuel pump to a pressure sensor. An aircraft fuel pump system comprises an aircraft fuel pump and a pressure sensor connected the aircraft fuel pump via a sensor line. The sensor line includes a reservoir located between the aircraft fuel pump and the pressure sensor. The reservoir acts to prevent liquid contacting the pressure switch when the fuel pump is not active.

Description:
This application claims priority to UK Patent Application No. 1301527.6 filed 29 Jan. 2013, the entire content of which is hereby incorporated by reference. 
     BACKGROUND OF THE INVENTION 
     The present invention concerns a device for use in an aircraft fuel pressure sensor line. More particularly, but not exclusively, this invention concerns a fuel pressure sensor line connecting an aircraft fuel pump to a pressure sensor. The invention also concerns a fuel pressure sensor line reservoir for use with a sensor line connecting an aircraft fuel pump to a pressure sensor. 
     Large aircraft, such as the Airbus A380, include several fuel tanks, with fuel being stored in a number of fuel tanks located in the wings of the aircraft. In order to move the fuel from a fuel tank into an engine, or to move fuel between different fuel tanks during flight, an aircraft fuel tank may be provided with fuel transfer pumps. In order to be able to detect whether or not a fuel transfer pump is working, a sensor line may be connected to a feed line leading from the fuel pump outlet, the sensor line leading to a pressure switch associated with the sensor line. The pressure switch may comprise a diaphragm and a certain amount of residual air. When the pump is operational, the feed line pushes the fuel and water mix typically found in an aircraft fuel tank up the sensor line towards the pressure switch compressing the air in the switch. The pressure increase due to the fuel flow pressure, a typical example of which is 30 psi, activates the pressure switch and provides an indication that the fuel pump is working correctly. If the fuel pressure switch is not activated, a monitoring system may inform the aircraft operator that the fuel pump is not working, for example using a warning light and/or audible alarm. 
     However, the above described operation may cause problems. As the fuel pump goes through operational cycles of being on and off, water may build up in the pressure switch, potentially due to the greater density of water compared to aircraft fuel. The pressure switch may begin to act as a sump and collect the water. The water in the switch may then freeze and prevent the proper operation of the pressure switch. This may result in an indication that the fuel pump is not operating correctly, even when it is operating correctly. An investigation of the fault may cause the aircraft to be grounded, thus increasing the downtime of the aircraft, causing an airline operator a financial loss. The problem of pressure switches freezing may be most noticeable in the outer wing fuel tanks, due to the low relative height of the pressure switches compared to the fuel pumps. As an aircraft takes off and pitches to 18 degrees, any residual fuel/water mix that has remained in the feed line and sensor line can flow into the pressure switch where it again acts as a sump and retains the unwanted water, thus leaving the pressure switch vulnerable to freezing. 
     Proposals for overcoming the problem of freezing pressure switches include the following. 
     One proposed solution is to position the fuel pump and pressure switch such that the sensor line maintains a positive gradient during take-off, thus preventing the fuel/water mix travelling into the pressure switch as the aircraft pitch increases. Such a solution may be used in fuel tanks inboard of the outer wing tanks and pressure switches in such tanks have a lower failure rate than when this is not the case. However, such arrangements may not be feasible in existing aircraft, because of the cost of retrofitting such a solution, or because of the space restrictions in those aircraft fuel tanks. 
     An alternative solution is to change the route of the sensor line such that it maintains a positive gradient during takeoff. However, similar design restrictions as listed above apply when attempting to reroute the sensor line. Also, such a reroute may require the addition of a sharp negative angle into part of the sensor line, which could act to channel any condensation or water build up into the pressure switch. 
     Another alternative solution is to fill the pressure switch with Ethylene Glycol, which acts as an anti-freeze when mixed with water. However, this is a temporary solution only as the fuel/water mix will still enter the pressure switch and could eventually flush out the Ethylene Glycol. 
     The present invention seeks to mitigate the above-mentioned problems. 
     SUMMARY OF THE INVENTION 
     The present invention provides, according to a first aspect, an aircraft fuel pump system comprising: 
     an aircraft fuel pump; and 
     a pressure sensor connected to the aircraft fuel pump via a sensor line, wherein the sensor line includes a reservoir located between the aircraft fuel pump and the pressure sensor. 
     The reservoir acts to hold any residual fuel/water mix away from the pressure switch when the fuel pump is not in operation. Preferably, the reservoir acts to define a wet sensor line between the aircraft fuel pump and the reservoir and a dry sensor line between the reservoir and pressure switch. The wet sensor line preferably contains a fuel/water mix. The dry sensor line preferably contains air, more preferably with no fuel/water mix in the dry sensor line. The reservoir is preferably arranged to contain any fluid pushed into the reservoir by the activation of the aircraft fuel pump, and transmit the pressure increase to the pressure switch by allowing the compression of the air present in the dry sensor line. By keeping the pressure switch as dry as possible, the chances of the switch freezing are reduced. 
     The reservoir may comprise an input, a sump, and an output. The reservoir may be configured such that there is no direct linear flow path between the input and the output. That there is no direct linear flow path between the input and the output means that fluid entering reservoir by the input, if maintaining the same direction of flow as taken through the input, will not directly flow into and out of the output. Instead, the fluid entering the reservoir via the input flow into the sump, where the flow direction of the fluid must change in order to leave the sump via the output. The reservoir may be configured such that the sump must be substantially full of a fuel/water mix before the fuel/water mix can pass through the reservoir. Advantageously, in such an arrangement, the flow of the fuel/water mix will first pass through the reservoir input, fill the sump, and only then pass through the reservoir output. Preferably, the reservoir contains air. Preferably, as a fuel/water mix enters the sump, air within the reservoir is compressed. Preferably, the compressed air transmits pressure via the output of the reservoir to the pressure switch. Preferably, the compressed air transmits pressure via the output of the reservoir and the dry sensor line to the pressure switch. Preferably, the reservoir is configured such that a fuel/water mix can only pass through the reservoir when substantially all of the air present in the reservoir has been expelled. Preferably, the sensor line and pressure sensor provides a closed circuit, i.e. fuel/water and air does not pass through the pressure sensor for recirculation back around into the aircraft fuel tanks. 
     Advantageously, the internal geometry of the reservoir inhibits either or both of the horizontal and vertical movement of the fuel/water mix during extreme aircraft pitch angles and/or negative G environments. Preferably, the internal geometry of the reservoir inhibits the flow of the fuel/water mix through the reservoir during extreme aircraft pitch angles and/or negative G environments. The reservoir may be configured to protect against pressure spikes during the fuel pump operation damaging the pressure sensor. 
     Experimental tests of fuel sensor lines according to the prior art have shown that the sensor line only fills to a volume of 70% during operation of the fuel pump, with the remaining volume being filled by air. The reservoir may be placed such that the fuel/water mix remains in the sensor line between the fuel pump and the reservoir, and the reservoir itself, with the sensor line between the pressure switch and the reservoir remaining mostly dry. The positioning and configuration of the reservoir to achieve this may be determined experimentally. The reservoir is preferably configured such that any fuel/water mix remaining in the sensor line between the fuel pump and the reservoir when the aircraft fuel pump is not operational is prevented from travelling past the reservoir towards the pressure switch due to the sump receiving the fuel/water mix. The configuration of the reservoir is preferably such that the air present between the reservoir and the pressure switch acts as an air bubble, preventing the flow of fuel/water mix between the reservoir and the pressure switch. 
     According to a second aspect of the invention there is also provided a reservoir for installation on a sensor line in an aircraft fuel pump system, the reservoir comprising an inlet, a sump, and an outlet, configured such that the sump must be substantially full before liquid can pass through the reservoir. 
     According to a third aspect of the invention, there is provided an aircraft fuel tank, the aircraft fuel comprising an aircraft fuel pump system as described above. 
     According to a fourth aspect of the invention, there is provided an aircraft wing, the aircraft wing comprising an aircraft fuel tank including an aircraft fuel pump system as described above. 
     According to a fifth aspect of the invention, there is provided an aircraft, the aircraft comprising a fuel tank including an aircraft fuel pump system as described above. 
     It will of course be appreciated that features described in relation to one aspect of the present invention may be incorporated into other aspects of the present invention. For example, the first aspect of the invention may incorporate any of the features described with reference to the second aspect of the invention and vice versa. 
    
    
     
       DESCRIPTION OF THE DRAWINGS 
       Embodiments of the present invention will now be described by way of example only with reference to the accompanying schematic drawings of which: 
         FIG. 1  shows a schematic view of an aircraft fuel pump system according to the prior art; 
         FIG. 2  shows a schematic view of an aircraft fuel pump system according to a first embodiment of the invention; 
         FIG. 3  shows an external view of a reservoir according to a second embodiment of the invention; 
         FIG. 4A  shows a cross-sectional view of a first possible internal configuration of the reservoir shown in  FIG. 3 ; 
         FIG. 4B  shows a cross-sectional view of a second possible internal configuration of the reservoir shown in  FIG. 3  and 
         FIG. 5  shows an external view of a reservoir according to a third embodiment of the invention; 
         FIG. 6A  shows a cross-sectional view of a first possible internal configuration of the reservoir shown in  FIG. 5 ; 
         FIG. 6B  shows a cross-sectional view of a second possible internal configuration of the reservoir shown in  FIG. 5 ; and 
         FIG. 7  shows an aircraft including an aircraft wing including an aircraft fuel tank according to the invention. 
     
    
    
     DETAILED DESCRIPTION 
       FIG. 1  shows a prior art fuel tank system, comprising an aircraft fuel tank  10  including an aircraft fuel pump  12 , the aircraft fuel pump  12  being connected to a pressure sensor  14  via a sensor line  16 . As has been explained, when the aircraft fuel pump  12  is in operation, the sensor line  16  passes a fuel/water mix towards the pressure sensor  14 , which is activated and may send a signal to an aircraft control unit  18  indicating that the aircraft fuel pump  12  is operating correctly. The sensor line  16  contains a fuel/water mix and a certain amount of air, and the fuel/water mix may enter the pressure switch, leaving it vulnerable to freezing. The aircraft fuel pump  12  is connected to a transfer pipe (not shown) which transmits fuel from one aircraft fuel tank to another. 
       FIG. 2  shows a first aspect of the present invention. An aircraft fuel tank  20  includes an aircraft fuel pump  22 , a reservoir  24 , and a pressure sensor  26 . The pressure sensor  26  is attached to a control unit as is conventional in the prior art. The aircraft fuel pump  22  is connected to a transfer pipe (not shown to improve clarity) as is conventional in the prior art. The aircraft fuel pump  22  is connected with the reservoir  24  via a first section of sensor line  28 , also known as the wet sensor line, and the reservoir  24  is also connected with the pressure sensor  26  via a second section of sensor line  30 , also known as the dry sensor line. The first section of sensor line  28  feeds into the reservoir  24  via an input (not shown in  FIG. 2 ) and the second section of sensor line  30  feeds out of the reservoir  24  via an output (not shown in  FIG. 2 ). As will be better illustrated in  FIGS. 3 to 6 , there is no direct linear flow path between the input and output of the reservoir  24 . 
       FIG. 3  shows an external view of a reservoir  40  according to one aspect of the invention, the reservoir comprising an input  42  and an output  44 . 
       FIG. 4A  shows a cross sectional view of a first possible internal configuration of the reservoir  40 . The input  42  feeds into a sump  46 , the sump  46  extending in the same direction as the fluid flow through the input  42 . The direction of flow through the input  42  is indicated by the arrow A. The output  44  is offset from the input  42 , and runs in a parallel direction to the input  42 . As can be seen in  FIG. 4A , the sump extends beyond the opening of the output  44  in what may be considered a “downstream” direction. Therefore, a fuel/water mix entering the reservoir via the input  42  will first travel to the end of the sump  46 , and fill the sump  46 , before being able to travel out of the reservoir via the output  44 . The reservoir is arranged such that when the fuel pump system is installed in an aircraft fuel tank, the input  42  is generally arranged to be oriented below the output  44  when the aircraft is at an approximately level pitch. During a flight, a change of pitch of the aircraft to which the system is installed may result in the input  42  being oriented above the output  44 , but the configuration of the reservoir  40  is such that the fuel/water mix should not pass beyond the reservoir  40  towards the pressure switch in the system. 
     As the aircraft fuel pump is activated, a fuel/water mix is transmitted into the reservoir  40  via the inlet  42 . Air present in the sump  46  will be displaced, compressing the air present in the outlet  44 , the compression of which goes on to activate a pressure switch. The amount of air present in the system preferably does not allow the fuel/water mix entering the sump  46  to pass through the reservoir  40 , due to the level of compression of the air being required being too great to be achieved by the aircraft fuel pump. 
       FIG. 4B  shows a cross-sectional view of a second possible internal configuration of the reservoir  40 . The outlet  44 ′ is configured in the same way as in  FIG. 4A , but the inlet  42 ′ extends into the sump  46 ′ as shown. As described above, air within the reservoir acts under compression so as to prevent the passage of fuel/water mix through the reservoir. 
       FIG. 5  shows an external view of a reservoir  50  according to a third aspect of the invention. The reservoir  50  is approximately cylindrical and includes an outlet  54 . The outlet  54  is associated with an expanded portion  58 , which expands the cross-section of the reservoir  50  beyond the circumference of the cylinder. The expanded portion  58  is arranged to increase the amount of air capable of being stored within the reservoir  50  in proximity to the outlet  54 . 
       FIGS. 6A and 6B  show a first possible and second possible internal configuration of the reservoir  50 . As the internal configurations correspond approximately to those shown in  FIGS. 4A and 4B , no great detail will be provided.  FIG. 6A  shows the reservoir  50  including an inlet  52 , and outlet  54 , and a sump  56 . The main difference with the embodiment shown in  FIG. 4A  is the presence of the expanded portion  58 , which acts to increase the air storage space within the reservoir  50  compared to a similarly dimensioned reservoir  40 .  FIG. 6B  shows the reservoir  50 ′ including an inlet  52 ′, and outlet  54 ′, and a sump  56 ′. In a preferred embodiment, during activation of an aircraft fuel pump the sump  56 ′ is configured to receive a fuel/water mix in the bottom part of the sump, immediately adjacent to the inlet  52 ′, and the top part of the sump  56 ′ is filled with air, as indicated in the figure. The embodiment shown in  FIG. 6B  can be seen to be configured to make it even more difficult for a fuel/water mix to pass through the reservoir and into a pressure switch. It can be seen, as for the embodiments described above also, that there is no direct fluid flow path between the input and output of the reservoir, and that in order to pass through the reservoir, fluid first has to substantially fill the sump  56 , which due to the amount of air within the reservoir is not usually possible when installed in an aircraft fuel system. 
       FIG. 7  shows an aircraft  70  including a wing  72 . The wing  72  includes an aircraft fuel tank  74 , with the aircraft fuel tank including a fuel pump sensor system according to any aspect of the invention described above. 
     Whilst the present invention has been described and illustrated with reference to particular embodiments, it will be appreciated by those of ordinary skill in the art that the invention lends itself to many different variations not specifically illustrated herein. 
     Where in the foregoing description, integers or elements are mentioned which have known, obvious or foreseeable equivalents, then such equivalents are herein incorporated as if individually set forth. Reference should be made to the claims for determining the true scope of the present invention, which should be construed so as to encompass any such equivalents. It will also be appreciated by the reader that integers or features of the invention that are described as preferable, advantageous, convenient or the like are optional and do not limit the scope of the independent claims. Moreover, it is to be understood that such optional integers or features, whilst of possible benefit in some embodiments of the invention, may not be desirable, and may therefore be absent, in other embodiments.