Aeronautical equipment

Abstract

This aeronautical equipment for an aircraft, comprising a part configured to be positioned at the level of a skin of the aircraft and means for reheating this part comprising a closed-circuit thermodynamic loop in which a phase-change heat transfer fluid circulates, is wherein it includes means for monitoring the fluid pressure in the loop in order to detect and report a malfunction of the equipment.

Claims

1. An aeronautical equipment for an aircraft, comprising a part configured to be positioned at the level of a skin of the aircraft and means for reheating this part comprising a closed-circuit thermodynamic loop in which a phase-change heat transfer fluid circulates, wherein the aeronautical equipment includes means for monitoring the fluid pressure in the loop in order to detect and report a malfunction of the equipment when the pressure in the loop drops and is at about ambient atmospheric pressure.

2. The aeronautical equipment for an aircraft according to claim 1, wherein the closed circuit further comprises at least one evaporator associated with heating means and at least one compensation chamber.

3. The aeronautical equipment for an aircraft according to claim 2, wherein the heating means comprise at least one heating resistance.

4. The aeronautical equipment for an aircraft according to claim 2, wherein the compensation chamber is associated with the monitoring means.

5. The aeronautical equipment for an aircraft according to claim 4, wherein the monitoring means comprise a pressure sensor.

6. The aeronautical equipment for an aircraft according to claim 5, wherein the pressure sensor is located in the compensation chamber.

7. The aeronautical equipment according to claim 1, wherein the aeronautical equipment comprises an aerodynamic measuring probe.

Description

BRIEF DESCRIPTION OF THE DRAWING

(1) The invention will be better understood using the following description, provided solely as an example and done in reference to the appended drawing, which shows a block diagram illustrating the structure of an aeronautical equipment for an aircraft, according to the invention.

(2) FIG. 1 This figure in fact illustrates an aeronautical equipment for an aircraft for example comprising an aerodynamic measuring probe designated by general reference 1.

DETAILED DESCRIPTION

(3) Such an aerodynamic measuring probe for example includes a Pitot tube.

(4) This Pitot tube then conventionally includes a part configured to be positioned at the level of a skin of the aircraft, and which is designated by general reference 2, and means for reheating this part, which are designated by general reference 3.

(5) In this illustrated exemplary embodiment, these reheating means 3 comprise a closed-circuit thermodynamic loop, with a base of fluid channels included in its structure and in which a phase-change heat transfer fluid circulates.

(6) Conventionally, this closed circuit further comprises at least one evaporator designated by general reference 4, associated with heating means designated by general reference 5 and at least one compensation chamber designated by general reference 6.

(7) The heating means for example comprise a heating resistance.

(8) The latter are electrically connected to the electric grid of the aircraft, which supplies the electrical power.

(9) Channels are provided in the rest of the equipment to constitute the thermodynamic circulation loop of the phase-change heat transfer fluid.

(10) According to the invention, the aeronautical equipment is also equipped with means for monitoring the fluid pressure in the loop in order to detect a malfunction thereof.

(11) These means for monitoring the fluid pressure in the loop are designated by general reference 7 in this figure and are associated with the compensation chamber 6.

(12) Indeed, these monitoring means for example comprise a pressure sensor that is for example placed in the compensation chamber 6, this sensor being designated by general reference 8.

(13) Of course, this sensor could be placed elsewhere in the fluid reheating circuit.

(14) This pressure sensor is then suitable for detecting and reporting a malfunction of the equipment when the pressure in the circuit drops and in particular is close to the atmospheric pressure.

(15) Indeed, the pressure variations of the system during normal operation are not comparable to a variation of an Off-Line system (variation in atmospheric pressure around 1 bar).

(16) Indeed, at 20° C., the saturation pressure for the fluid experienced in the application yields a value of 8.5 bars. Above 70 to 100° C., the pressure is between 39 and 89 bars, and these pressure variations are therefore easily detectable.

(17) This pressure sensor then acts to open a circuit connecting the electrical power of this equipment to the rest of the aircraft for example by opening this circuit by means of a switch 9.

(18) Thus, such monitoring means make it possible to detect a failure of the diphasic system by opening a contact after verification of the pressure in the compensation chamber.

(19) One thus reproduces the operation equivalent to a cut heating wire like in the current systems, which makes it possible to use the aeronautical equipment according to the invention with current aircraft without modifying airplane interfaces and without adding surveillance electronics.

(20) Of course, still other embodiments can be considered.