阅读说明：本技术 用于使显示系统对故障的灵敏度降低的装置和方法 (Apparatus and method for reducing sensitivity of display system to failure ) 是由 J·古里卢 A·佩兰 于 2019-05-20 设计创作，主要内容包括：本发明涉及用于使显示系统对故障的灵敏度降低的装置和方法。减敏装置(1)包括：获取模块(3),获取模块用于获取飞行器的当前飞行参数；生成模块(4),所述生成模块用于生成描绘飞行符号体系的第一图像,所述飞行符号体系包括至少一个零俯仰参考线；生成模块(5),生成模块用于生成描绘所述飞行器所飞越的地形的地貌的第二图像；生成模块(6),生成模块用于生成包括由分割线分开的至少两个部分的第三图像(I3)；生成模块(7),生成模块被配置为通过依次叠加所述第一图像、所述第二图像和所述第三图像来生成第四图像,所述第一图像放置在前景中；以及传输模块(8),传输模块被配置为向用户装置(9)传输表示所述第四图像的信号。(the present invention relates to an apparatus and method for reducing the sensitivity of a display system to faults. The desensitization device (1) comprises: the acquisition module (3) is used for acquiring the current flight parameters of the aircraft; a generation module (4) for generating a first image depicting a flight symbology, the flight symbology comprising at least one zero-pitch reference line; a generation module (5) for generating a second image depicting the topography of the terrain over which the aircraft is flying; a generating module (6) for generating a third image (I3) comprising at least two parts separated by a dividing line; a generation module (7) configured to generate a fourth image by superimposing in sequence the first image, the second image and the third image, the first image being placed in a foreground; and a transmission module (8) configured to transmit a signal representing the fourth image to a user device (9).)

1. a method for desensitizing an Aircraft (AC) display system configured to display a composite representation of terrain to faults,

Characterized in that the method comprises the following steps:

-an acquisition step (E1) carried out by an acquisition module (3) comprising acquisition of environmental data and current flight parameters of the Aircraft (AC), said current flight parameters comprising at least attitude, altitude, three-dimensional position and heading;

-a first generation step (E2) carried out by a first generation module (4) comprising generating a first image (I1) depicting a flight symbology comprising at least one zero-pitch reference line, the first image (I1) indicating a pitch angle with respect to a zero pitch angle of the Aircraft (AC), the zero-pitch reference line representing the zero pitch angle of the Aircraft (AC), the first image (I1) being generated at least on the basis of the attitude;

-a second generation step (E3), carried out by a second generation module (5), comprising generating a second image (I2) depicting the topography of the terrain over which the Aircraft (AC) flies, based at least on the attitude, the heading, the three-dimensional position and the environmental data;

-a third generation step (E4) carried out by a third generation module (6) comprising generating a third image (I3) comprising at least two portions separated by a dividing line determined on the basis of the attitude of the Aircraft (AC), a first portion (13) comprising pixels having values representative of a first range of colors and a second portion (14) comprising pixels having values representative of a second range of colors;

-a fourth generation step (E5), carried out by a fourth generation module (7), comprising generating a fourth image (I4) by superimposing in sequence the first image (I1), the second image (I2) and the third image (I3), the first image (I1) being placed in the foreground such that, when the topography of the flying terrain shown in the second image shifts downwards due to a fault, the second portion of the third image is visible in the fourth image, highlighting the fault;

-a transmission step (E6), carried out by a transmission module (8), comprising the transmission of a signal representative of the fourth image (I4) to a user device (9).

2. The method of claim 1, wherein the first and second light sources are selected from the group consisting of,

characterized in that said second generation step (E3) comprises the following sub-steps:

-a first determining sub-step (E31), carried out by a first determining sub-module (51), comprising determining a topography of the terrain over which the Aircraft (AC) flies, based at least on the attitude, the three-dimensional position, the heading and the environmental data;

-a first assigning sub-step (E32), implemented by a first assigning sub-module (52), comprising assigning at least one pixel value to a first set of pixels forming part of the second image (I2), said one or more pixel values representing the topography of the terrain over which the Aircraft (AC) flies;

-a second assigning sub-step (E33), implemented by a second assigning sub-module (53), comprising assigning values representative of transparent pixels to a second set of pixels forming part of the second image (I2), corresponding to a set of pixels of the second image (I2) whose pixels do not belong to the first set of pixels.

3. The method according to any one of claims 1 and 2,

characterized in that said third generation step (E4) comprises the following sub-steps:

-a second determining sub-step (E41) carried out by a second determining sub-module (61) comprising determining the position of the dividing line based on the zero pitch reference line, the dividing line being parallel to the zero pitch reference line (10), the second portion (14) of the third image (I3) corresponding at least in part to the topography of the terrain (15) flown over by the Aircraft (AC) represented in the second image (I2) when the first image (I1), the second image (I2) and the third image (I3) are superimposed in the fourth generating step (E5);

-an adding sub-step (E42), implemented by an adding module (62), comprising adding an indicative element (12) indicative of an error to the second portion (14) of the third image (I3) in order to unambiguously signal the fault, this indicative element (12) being visible in the fourth image when the topography of the flying terrain shown in the second image is shifted downwards due to the fault.

4. The method of claim 3, wherein the first and second light sources are selected from the group consisting of,

Characterized in that the dividing line (11) is arranged in a position corresponding to a zero pitch angle of the Aircraft (AC);

And in that the indicative element (12) is arranged at a distance with respect to the division line (11) corresponding to a pitch angle (Δ ⁰) corresponding to an angle defined by the difference between the zero pitch angle and the natural horizon of the Aircraft (AC) when the Aircraft (AC) is flying at maximum altitude.

5. The method of claim 3, wherein the first and second light sources are selected from the group consisting of,

Characterized in that the division line (11) is arranged at a pitch angle (Δ ⁰) relative to a zero pitch angle position of the Aircraft (AC) equal to the difference between the zero pitch angle of the Aircraft (AC) and the natural horizon when the Aircraft (AC) is flying at maximum altitude.

6. the method of claim 3, wherein the first and second light sources are selected from the group consisting of,

characterized in that the division line (11) is arranged at a pitch angle (Δ ⁰) in relation to a zero pitch position of the Aircraft (AC), which pitch angle depends on the altitude of the Aircraft (AC).

7. The method of claim 3, wherein the first and second light sources are selected from the group consisting of,

Characterized in that the thickness (E) of the dividing line (11) is equal to a certain pitch angle (Δ ⁰) defined by the difference between a zero pitch angle of the Aircraft (AC) when the Aircraft (AC) is flying at maximum altitude and a natural horizon, the dividing line (11) comprising pixels exhibiting a gradient of values along the thickness (E) between a value equal to a first range of colors representing a first row of pixels adjacent to the first portion (13) of the third image (I3) and a value equal to a second range of colors representing a second row of pixels adjacent to the second portion (14) of the third image (I3), the first row of pixels being arranged in a position corresponding to the zero pitch angle of the Aircraft (AC).

8. The method of claim 3, wherein the first and second light sources are selected from the group consisting of,

Characterized in that the division line (11) is arranged at a pitch angle (Δ ⁰) relative to a zero pitch angle position of the Aircraft (AC), which pitch angle is equal to the difference between the zero pitch angle of the Aircraft (AC) and the horizon, which horizon is determined on the basis of a predetermined distance (17) defined by a maximum field of view (21) of the topography of the terrain (15) represented in the second image (I2) when the Aircraft (AC) is flying at maximum altitude.

9. the method of claim 3, wherein the first and second light sources are selected from the group consisting of,

characterized in that the division line (11) is arranged at a pitch angle (Δ ⁰) with respect to a zero pitch position of the Aircraft (AC), the pitch angle being dependent on the altitude of the Aircraft (AC) and on a predetermined distance (17) defined by a maximum field of view (21) of the topography of the terrain (15) represented in the second image (I2).

10. the method of claim 3, wherein the first and second light sources are selected from the group consisting of,

Characterized in that the thickness (E') of the dividing line (11) is equal to a certain pitch angle (Δ ⁰) defined by the difference between the natural horizon and the horizon depending on a predetermined distance (17) defined by the maximum field of view (21) of the topography of the terrain (15) represented in the second image (I2), the dividing line (11) comprising pixels exhibiting values representing a third color range, the dividing line (11) being arranged at a pitch angle relative to a zero pitch position of the Aircraft (AC), the pitch angle being equal to the difference between the zero pitch angle and the natural horizon.

11. the method according to claim 3 or 10,

characterized in that the division line (11) is arranged at a pitch angle (Δ ⁰) with respect to a zero pitch position of the Aircraft (AC), said pitch angle being dependent on the altitude of the Aircraft (AC) minus the altitude of the terrain over which the Aircraft (AC) is flying and on a predetermined distance (17) defined by a maximum field of view (21) of the topography of the terrain represented in the second image (I2).

12. The method of claim 11, wherein the first and second light sources are selected from the group consisting of,

characterized in that if the Aircraft (AC) itself is in one of the following orientations:

A final approach in the landing phase,

During the takeoff phase

-in the phase of the missed approach,

The height of the terrain flown over is equal to the current height of the terrain flown over in one of the preceding cases,

Otherwise, the altitude of the terrain in flight is assumed to be zero.

13. The method of claim 11, wherein the first and second light sources are selected from the group consisting of,

Characterized in that the height of the terrain flown over is equal to the minimum height of the terrain flown over by the Aircraft (AC) within a section of a cylinder centred on the Aircraft (AC), the radius of the section of cylinder (20) being equal to the predetermined distance (17) and having the aperture angle (θ) of the maximum field of view (21).

14. An apparatus for desensitizing an Aircraft (AC) display system configured to display a composite representation of terrain to faults,

The device is characterized by comprising the following modules:

-an acquisition module (3) configured to acquire environmental data and current flight parameters of the Aircraft (AC), said current flight parameters comprising at least an attitude, an altitude, a three-dimensional position and a heading;

-a first generation module (4) configured to generate a first image (I1) depicting a flight symbology, the flight symbology comprising at least one zero-pitch reference line, the first image (I1) indicating a pitch angle with respect to a zero pitch angle of the Aircraft (AC), the zero-pitch reference line representing the zero pitch angle of the Aircraft (AC), the first image (I1) being generated based at least on the attitude;

-a second generation module (5) configured to generate a second image (I2) depicting a topography of the terrain over which the Aircraft (AC) flies, based on at least the attitude, the heading, the three-dimensional position and the environmental data;

-a third generation module (6) configured to generate a third image (I3) comprising at least two parts separated by a dividing line determined based on the pose of the Aircraft (AC), a first part comprising pixels having values representing a first color range and a second part comprising pixels having values representing a second color range;

-a fourth generation module (7) configured to generate a fourth image (I4) by superimposing in sequence the first image (I1), the second image (I2) and the third image (I3), the first image (I1) being placed in the foreground such that the second portion of the third image is visible in the fourth image, highlighting the fault, when the topography of the flown-over terrain shown in the second image shifts downwards due to the fault;

-a transmission module (8) configured to transmit a signal representative of the fourth image (I4) to a user device (9).

15. An aircraft is provided, which is provided with a plurality of flying wheels,

Characterized in that the aircraft comprises a fault desensitizing device according to claim 14.

Technical Field

The present invention relates to an apparatus and method for reducing the sensitivity to faults of a display system, particularly a primary flight display configured to display a composite topography as a background image for flight and navigation symbology.

Background

the Primary Flight Display (PFD) typically displays images depicting flight and navigation symbologies generated by a Control and Display System (CDS) or by an Electronic Instrumentation System (EIS). In parallel, a Synthetic Vision System (SVS) generates an image depicting the topology of the terrain flown over by the aircraft. This image, most recently, is placed in the context of an image depicting a flight and navigation symbology. The display may also be disposed on a display surface such as a head-up display (HUD). This surface corresponds to a transparent surface in the external field of view of the pilot that displays flight information.

This display enhances the situational awareness of the pilot with respect to the aircraft's geographic location, the aircraft's location relative to the terrain over which it is flying, and the aircraft's relative location with respect to the approach trajectory for landing.

however, this display may fail if pitch data or altitude data is corrupted (changed), for example, during processing by the SVS. These faults may cause the topology of the terrain in flight in the display to shift downward or upward.

Disclosure of Invention

the object of the present invention is to overcome these drawbacks by providing a fault desensitization method and device while limiting the pilot's load if the topology shifts.

to this end, the invention relates to a method for desensitizing an aircraft display system configured to display a composite representation of terrain to faults.

According to the invention, the desensitization method comprises the following steps:

-an acquisition step, performed by an acquisition module, comprising acquiring environmental data and current flight parameters of the aircraft, said current flight parameters comprising at least an attitude, an altitude, a three-dimensional position and a heading;

-a first generation step, performed by a first generation module, comprising generating a first image depicting a flight symbology, the flight symbology comprising at least one zero-pitch reference line, the first image indicating a pitch angle relative to a zero-pitch angle of the aerial vehicle, the zero-pitch reference line representing the zero-pitch angle of the aerial vehicle, the first image being generated based at least on the attitude;

-a second generation step, performed by a second generation module, comprising generating a second image, based on at least the attitude, the heading, the three-dimensional position and the environmental data, the second image depicting a topography of a terrain over which the aircraft is flying;

-a third generation step, carried out by a third generation module, comprising generating a third image comprising at least two portions separated by a segmentation line determined based on the attitude of the aircraft, a first portion comprising pixels having values representative of a first color range and a second portion comprising pixels having values representative of a second color range;

-a fourth generation step, performed by a fourth generation module, comprising generating a fourth image by superimposing in sequence said first image, said second image and said third image, said first image being placed in the foreground such that when the topography of the overflowed terrain, shown in said second image, is displaced downwards due to a fault, said second portion of said third image is visible in said fourth image, highlighting said fault;

-a transmission step, implemented by a transmission module, comprising the transmission of a signal representative of said fourth image to a user device.

Thus, by means of the third image generated by the third generating module, a fault can be highlighted by the second part of said third image appearing on the display when the topography of the terrain flown is shifted downwards.

according to one feature, the second generation step comprises the sub-steps of:

-a first determining sub-step, implemented by a first determining sub-module, comprising determining a topography of the terrain over which the aircraft is flying, based at least on the attitude, the three-dimensional position, the heading, and the environmental data;

-a first assigning sub-step, implemented by a first assigning sub-module, comprising assigning at least one pixel value to a first set of pixels forming part of the second image, the one or more pixel values representing the topography of the terrain over which the aircraft is flying;

-a second assigning sub-step, implemented by a second assigning sub-module, comprising assigning values representative of transparent pixels to a second set of pixels forming part of the second image, the second set of pixels corresponding to a set of pixels of the second image whose pixels do not belong to the first set of pixels.

According to another feature, the third generation step comprises the following sub-steps:

-a second determining sub-step, implemented by a second determining sub-module, comprising determining the position of the dividing line based on the zero-pitch reference line, the dividing line being parallel to the zero-pitch reference line, the second portion of the third image corresponding at least in part to the topography of the terrain over which the aircraft is flying, represented in the second image, when the first image, the second image and the third image are superimposed in the fourth generating step;

-an adding sub-step, implemented by an adding module, comprising adding an indicative element indicative of an error to the second portion of the third image, in order to explicitly signal the fault, this indicative element being visible in the fourth image when the topography of the terrain in flight shown in the second image is shifted downwards due to the fault.

according to a first embodiment, the division line is arranged in a position corresponding to a zero pitch angle of the aircraft;

and the indicative element is arranged at a distance relative to the dividing line corresponding to a pitch angle corresponding to an angle defined by the difference between the zero pitch angle of the aircraft and the natural horizon when the aircraft is flying at maximum altitude.

According to a second embodiment, the division line is arranged at a pitch angle relative to a zero pitch angle position of the aircraft, the pitch angle being equal to the difference between the zero pitch angle of the aircraft and the natural horizon when the aircraft is flying at maximum altitude.

According to a third embodiment, the division line is arranged at a pitch angle with respect to a zero pitch position of the aircraft, the pitch angle being dependent on the altitude of the aircraft.

according to a fourth embodiment, the thickness of the dividing line is equal to a pitch angle defined by the difference between zero pitch angle of the aircraft when the aircraft is flying at maximum altitude and the natural horizon, the dividing line comprising pixels exhibiting a gradient of values along the thickness between a value equal to a first range of colors representing a first row of pixels adjacent to the first portion of the third image and a value equal to a second range of colors representing a second row of pixels adjacent to the second portion of the third image, the first row of pixels being arranged in a position corresponding to the zero pitch angle of the aircraft.

According to a fifth embodiment, the division line is arranged at a pitch angle relative to a zero pitch angle position of the aircraft, the pitch angle being equal to the difference between the zero pitch angle of the aircraft and the horizon, the horizon being determined based on a predetermined distance defined by a maximum field of view of the topography of the terrain represented in the second image when the aircraft is flying at maximum altitude.

According to a sixth embodiment, the division line is arranged at a pitch angle with respect to a zero pitch position of the aircraft, the pitch angle being dependent on the altitude of the aircraft and on a predetermined distance defined by a maximum field of view of the topography of the terrain represented in the second image.

according to a seventh embodiment, the thickness of the dividing line is equal to a certain pitch angle defined by the difference between a natural horizon and a horizon depending on a predetermined distance defined by the maximum field of view of the topography of the terrain represented in the second image, the dividing line comprising pixels exhibiting values representing a third color range, the dividing line being arranged at a certain pitch angle with respect to a zero pitch position of the aircraft, the pitch angle being equal to the difference between the zero pitch angle and the natural horizon.

according to an eighth embodiment, the division line is arranged at a pitch angle with respect to a zero pitch position of the aircraft, the pitch angle being dependent on the altitude of the aircraft minus the altitude of the terrain over which the aircraft is flying and on a predetermined distance defined by a maximum field of view of the topography of the terrain represented in the second image.

Additionally, if the aircraft itself is in one of the following orientations:

A final approach in the landing phase,

during the takeoff phase

-in the phase of the missed approach,

The height of the terrain flown over is equal to the current height of the terrain flown over in one of the preceding cases,

Otherwise, the altitude of the terrain in flight is assumed to be zero.

According to a ninth and tenth embodiment, the altitude of the terrain flown over is equal to the minimum altitude of the terrain flown over by the aircraft within a cylindrical section, the cylinder being centered on the aircraft, the cylindrical section having a radius equal to the predetermined distance and having an aperture angle of maximum field of view.

the invention also relates to an apparatus for desensitizing an Aircraft (AC) display system configured to display a composite representation of terrain to faults.

according to the invention, the device comprises the following modules:

-an acquisition module configured to acquire environmental data and current flight parameters of the aircraft, the current flight parameters comprising at least an attitude, an altitude, a three-dimensional position and a heading;

-a first generation module configured to generate a first image depicting a flight symbology, the flight symbology comprising at least one zero-pitch reference line, the first image indicating a pitch angle relative to a zero-pitch angle of the aerial vehicle, the zero-pitch reference line representing the zero-pitch angle of the aerial vehicle, the first image being generated based on at least the attitude and the altitude;

-a second generation module configured to generate a second image based on at least the attitude, the heading, the three-dimensional position, and the environmental data, the second image depicting a topography of terrain over which the aircraft is flying;

-a third generation module configured to generate a third image comprising at least two parts separated by a segmentation line determined based on the pose of the aircraft, a first part comprising pixels having values representing a first color range and a second part comprising pixels having values representing a second color range;

-a fourth generation module configured to generate a fourth image by superimposing in sequence the first, second and third images, the first image being placed in the foreground such that when the topography of the terrain overflowed shown in the second image is displaced downwards due to a fault, the second portion of the third image is visible in the fourth image, thereby highlighting the fault;

-a transmission module configured to transmit a signal representing the fourth image to a user device.

the invention also relates to an aircraft, in particular a transport aircraft, comprising a fault-mitigation device as described above.

drawings

The invention and its features and advantages will become more apparent upon reading the description provided with reference to the drawings, in which:

Figure 1 schematically shows an embodiment of a fault desensitization device;

FIG. 2 schematically shows an embodiment of a fault desensitization method;

Fig. 3 shows an image displayed on the display device after the first image, the second image and the third image are superimposed;

Figure 4 shows a side view of an aircraft flying at low altitude, and a display comprising a first image and a second image in a superimposed state;

Figure 5 shows a side view of an aircraft flying at high altitude, and a display comprising a first image and a second image in a superimposed state;

fig. 6 shows a side view of an aircraft whose visibility of the SVS is limited, and a display comprising a first image and a second image in a superimposed state;

figure 7 shows a side view of an aircraft flying on terrain presenting a non-zero altitude and whose SVS is visibility-limited, and a display comprising a first image and a second image in a superimposed state;

fig. 8 shows a first image displaying a symbology on the left side and a third image on the right side according to some embodiments;

fig. 9 shows a first image displaying a symbology on the left and a third image on the right according to another embodiment;

fig. 10 shows a first image displaying a symbology on the left and a third image on the right according to another embodiment;

Fig. 11 shows a first image displaying a symbology on the left and a third image on the right according to another embodiment;

fig. 12 shows a fourth image when a fault occurs;

FIG. 13 shows a perspective view of an aircraft flying on a terrain presenting a non-zero altitude, and a cylindrical section representing the field of view of the SVS;

figure 14 shows a cross-sectional view of a cylinder segment divided into subsections;

FIG. 15 shows a side view of an aircraft flying on a terrain of non-zero altitude, and locations along a predetermined distance of maximum field of view for which an apparent angle of terrain (angle de terrain accommodation) has been determined;

Figure 16 shows a perspective view of an aircraft flying on a terrain of non-zero altitude, with the aim of determining the apparent angle of the terrain for a segment of a cylinder and a certain point;

fig. 17 shows on the left a first image showing the topographical angle of a subsection and on the right a third image generated on the basis of the smallest topographical angle;

Figure 18 shows the positioning of the indicative elements according to a variant;

Fig. 19 shows the positioning of the indicative elements according to another variant.

Detailed Description

Fig. 2 shows an embodiment of an apparatus 1 for desensitizing a display system (DISP)9 of an aircraft AC to faults. In the remainder of the description, the device 1 for desensitizing the display system to faults will be referred to as desensitizing device.

The display system 9 is configured to display a composite representation of the terrain. The display system 9 may correspond to a primary flight display PFD or a head-up display HUD surface.

the desensitization device 1 loaded on the aircraft AC comprises an acquisition module ACQ (also referred to as "module for acquisition") 3 configured to acquire current flight parameters and environmental data of the aircraft AC.

The acquisition module 3 acquires flight parameters measured by the sensors SENS1, SENS2, SENS3, SENS 4. The acquisition module 3 acquires at least one current attitude of the aircraft AC, a current altitude of the aircraft, a current three-dimensional position of the aircraft, a current heading of the aircraft AC, and current environmental data. The environmental data corresponds, for example, to data relating to the terrain overflown, to data relating to one or more runways, or to data relating to obstacles that may be encountered by the aircraft AC.

The desensitizing device 1 also comprises a generation module 4GEN2 (also referred to as "module for generation") configured to generate an image I1 depicting a flight symbology. The flight symbology includes at least one zero-pitch reference line 10. This second image I1 indicates a pitch angle relative to zero pitch angle of the aircraft AC. The zero pitch reference line 10 represents the zero pitch angle of the aircraft AC. Thus, the distance between the horizon and the zero-pitch reference line 10 represents the pitch angle relative to zero pitch angle of the aircraft AC. The image I1 is generated based on at least the current flight parameters corresponding to the current attitude of the aircraft AC. Image I1 includes a set of pixels. The pixels representing the symbology have values representing the colors displayed in image I1. The other pixels in the set have values representing transparent (colorless) pixels. The generation module 4 may be included in the CDS/EIS.

The desensitization device 1 further comprises a generation module GEN 15 configured to generate an image I2 depicting the topography of the terrain 15 flown over by the aircraft AC. The topography is generated based on current flight parameters corresponding to at least the attitude of the aircraft AC, the heading of the aircraft AC, the three-dimensional position of the aircraft AC, and the environmental data. The environmental data allows knowledge of the individual elements included within the topography of the terrain 15 being flown over. The attitude, heading, and three-dimensional position allow the topography determined based on the environmental data to be shown from the pilot's perspective. Thus, image I2 depicts the topography and sky of the terrain being flown over. The topography of the terrain being flown over and the sky are separated by the horizon 18.

For example, for a primary flight display PFD, the sky is represented by a blue color set and the topography of the terrain is represented by a green, brown, and/or white color set.

for example, for a head-up display HUD surface, the sky is transparent and the topography of the terrain is monochromatic. The topography of the terrain may be green in color.

Thus, when image I1 is superimposed on image I2, the topography is shown in the background relative to the symbology.

The desensitizing device 1 further comprises a generation module GEN 36 configured to generate an image I3 comprising at least two portions separated by a division line 11, the division line being determined according to the attitude of the aircraft AC, in particular according to the roll and pitch angles of the aircraft AC. Portion 13 comprises pixels having values representing a first color range and portion 14 comprises pixels having values representing a second color range.

the color range may correspond to a single color or a set of colors.

For example, for the primary flight display PFD, the first color range corresponds to a blue color, such that the color of the portion 13 is close to the color of the sky. The second color range corresponds to brown colors such that the color of the portion 14 approximates the color of the topography of the terrain.

according to another example, for a primary flight display PFD, the first color range corresponds to a set of colors representing the sky, such as a blue gradient. The second color range corresponds to a set of colors, such as a brown set of colors, that represent the topography of the terrain.

According to one example, for a heads-up display surface, the second color range corresponds to a set of green colors.

the generation module GEN 47 forming part of the desensitization device 1 is configured to generate an image I4 (fig. 3) by superimposing in sequence the image I1, the image I2 and the image I3. The image I1 is located in the foreground. Advantageously, the images I1, I2, I3 and I4 have the same pixel size, so that they can be directly superimposed.

A transmission module TRANS (also referred to as "module for transmission") 8 forming part of the desensitizing device 1 is configured to transmit signals representing the image I4 to a user device 9, such as a display system.

advantageously, the generation module 5 comprises a determination submodule DET1 (also called "module for determination") 51 configured to determine the topography of the terrain flown over by the aircraft AC at least on the basis of the attitude, the three-dimensional position, the heading and the environmental data.

assignment submodule ASSIGN1 (also referred to as "module for assigning") 52, which forms part of generation module 5, is configured to ASSIGN at least one pixel value to a set of pixels forming part of image I2 (fig. 3). The one or more pixel values assigned to the set of pixels represent the topography of terrain 15 flown over by the aircraft AC. The first color range of portion 13 of image I3 corresponds to one or more colors that are close to the one or more colors of the topography of image I2.

An assigning sub-module ASSIGN 253 forming part of the desensitizing arrangement 1 is configured to ASSIGN values representing transparent pixels to the sets of pixels 16 forming part of the image I2. The set of pixels 16 corresponds to a set of pixels from the image I2 whose pixels do not belong to the set of pixels 15.

Thus, by virtue of the transparency of these pixels of the set of pixels 16, the image I3 may appear in the background relative to the superimposed images I1 and I2. By virtue of the transparency of the pixels of the set of pixels 16, only the portion of the image I3 corresponding to the set of pixels 16 from the image I2 appears. At least a part of the portion 13 and possibly the portion 14 is present depending on the size of the set of pixels 16.

The generation module 6 may comprise a determination submodule DET 261 configured to determine the position of the dividing line 11 on the basis of the attitude of the aircraft AC. When all images I1, I2, and I3 have been superimposed, the dividing line 11 is parallel to the zero-pitch reference line 10. When image I1, image I2, and image I3 are superimposed by the generation module 7, the portion 14 of image I3 corresponds at least in part to the topography depicted in the second image I2 as being flown over the terrain 15 by the aircraft AC.

According to a variant embodiment, the determination submodule 61 may be included in the CDS/EIS.

According to another variant embodiment, the determination submodule 61 may be included in a system that exhibits comparable integrity in addition to the CDS and EIS.

an adding module ADD (also called "module for adding") 62, which forms part of the generating module 6, is configured to ADD an indicative element 12 indicating an error to the portion 14 of the image I3. The indicative element may correspond to an abbreviation, a graphic, text, or any other element representing the error. For example, the indicative element corresponds to text such as "SVS fail".

Thus, when the topology of the terrain 15 being flown over on the display shifts down due to data corruption, an indicative element 12 (fig. 12) appears which indicates that the generation module 5 is handling a fault in certain data, such as pitch and altitude data. When the topology of the flying terrain 15 is not shifted, the indicative element 12 is hidden by the topology of the flying terrain 15 in the image I2.

fig. 4 shows an image I5 displayed in an aircraft AC flying over terrain at low altitude, and in a display system without a desensitizing device 1 intended for this aircraft AC. Fig. 5 shows an aircraft AC flying over terrain at high altitude, and an image I5 intended to be displayed in the display system of the aircraft AC. Due to the curvature of the earth T, if the dividing line 11 is placed in a position corresponding to zero pitch angle, the indicative element 12 may appear even without a fault.

the following four embodiments provide solutions for dealing with the curvature of the earth T.

According to a first embodiment (fig. 8), the division line 11 is arranged in a position corresponding to a zero pitch angle of the aircraft AC, however, the indicative element 12 is arranged with respect to the division line 11 at a distance corresponding to a pitch angle Δ ⁰ corresponding to an angle defined by the difference between the zero pitch angle of the aircraft AC when the aircraft AC is flying at maximum altitude and the natural horizon.

For example, the following equation allows the pitch angle Δ ⁰ to be obtained from the altitude of the aircraft AC while taking into account the curvature of the earth T:

wherein, height_Footis the altitude (in feet) of the aircraft AC (1 foot equals approximately 0.305 m).

Thus, if the maximum altitude of the aircraft AC is assumed to be 50000 feet (about 15240m), the pitch angle is equal to Δ ⁰ equal to 1.71 °. thus, the indicative element 12 is located at a distance below the division line 11 corresponding to a pitch angle Δ ⁰ of 1.71 °.

according to the second embodiment, the division line 11 is arranged at a pitch angle Δ ⁰ with respect to the zero pitch position of the aircraft AC, the pitch angle Δ ⁰ is equal to the difference between the zero pitch angle of the Aircraft (AC) and the natural horizon when the Aircraft (AC) is flying at maximum altitude of the aircraft AC, therefore, the pitch angle Δ ⁰ is such that the division line 11 shown is located at the position of the natural horizon when the aircraft AC is flying at maximum altitude.

In this embodiment, the indicative element 12 may be "fixed" below the parting line 11. For example, the indicative element 12 may be placed at a distance between one and five pixels away from the dividing line 11.

According to the third embodiment, the division line 11 is arranged at a pitch angle Δ ⁰ with respect to the zero pitch position of the aircraft AC, which depends on the altitude of the aircraft AC.

In this embodiment, the position of the division line 11 is dynamic and varies with the altitude of the aircraft AC. The equation for determining the position of the dividing line may be the same as the equation described in the first embodiment, in which the altitude of the aircraft AC_FootIs variable. Other equations may be used. For example, other equations may consider the radius of the earth T or atmospheric diffraction effects. The indicative element 12 may be "fixed" below the parting line 11. For example, the indicative element 12 may be placed at a distance between one and five pixels away from the dividing line 11.

according to a fourth embodiment (fig. 10), the thickness E of the dividing line 11 is equal to the pitch angle defined by the difference between zero pitch angle of the aircraft AC when the aircraft AC is flying at maximum altitude and the natural horizon. Segmentation line 11 comprises pixels exhibiting a gradient of values along thickness E between a value equal to a first range of colors representing a first row of pixels adjacent to portion 13 of image I3 and a value equal to a second range of colors representing a second row of pixels adjacent to portion 14 of image I3. The first row of pixels is arranged in a position corresponding to a zero pitch angle of the aircraft AC.

It may also be advantageous to consider the limitations of the field of view 21 (fig. 13) caused by SVS in addition to the curvature of the earth T. Fig. 6 shows an aircraft AC flying over terrain, and an image I5 intended to be displayed in the display system of the aircraft AC. The horizon 18 of the feature 15 is shifted downward because of the limitation of the field of view 21 caused by the SVS. Thus, the indicative elements 12 tend to be exposed, since even without failure, the displayed topography of the terrain 15 is too low to hide the indicative elements 12.

The following three embodiments provide solutions for dealing with the curvature of the earth T and the limitations of the field of view 21 of the SVS.

According to a fifth embodiment (fig. 9), the division line 11 is arranged at a pitch angle Δ ⁰ with respect to a zero pitch angle position of the aircraft AC, which pitch angle is equal to the difference between the zero pitch angle of the aircraft AC and the horizon.

in this embodiment, the indicative element 12 may be "fixed" below the parting line 11. For example, the indicative element 12 may be placed at a distance between one and five pixels away from the dividing line 11.

for example, the following equation allows obtaining the pitch angle Δ ⁰ from the altitude of the aircraft AC while taking into account the curvature of the earth T and the limits of the field of view 21:

Wherein:

-height_FootIs the altitude (in feet) of the aircraft AC (1 foot is equal to about 0.305 m);

The SVS range _NM is a predetermined distance 17 (in nautical miles);

-A, B and C are constants.

Without limitation, constant a equals 0.3048, constant B equals 1852, and constant C equals 1.23.

Thus, the equation corresponds to a piecewise defined function. For heights above a maximum predetermined heightHeight of_footangle of pitch Δ⁰dependent on the altitude of the aircraft AC_footand a predetermined distance 17. To pairBelow a maximum predetermined heightHeight of (a), pitch angle delta⁰dependent only on the altitude of the aircraft AC_Foot。

thus, if it is assumed that the maximum altitude of the aircraft AC is equal to 50000 feet (about 15240m) and the predetermined distance 17 is equal to 50 nautical miles (about 92.6km), the pitch angle Δ ⁰ is equal to 9.35 °. the division line 11 is positioned at a pitch angle Δ ⁰ of 9.35 ° with respect to zero pitch angle.

according to a sixth embodiment (fig. 9), the division line is arranged at a pitch angle Δ ₀ with respect to the zero pitch position of the aircraft AC, which depends on the altitude of the aircraft AC and on the predetermined distance 17.

In this embodiment, the position of the division line 11 is dynamic and varies with the altitude of the aircraft AC. The equation for determining the position of the dividing line may be the same as the equation described in the fifth embodiment, in which the altitude of the aircraft AC_Footis variable. Other equations may be used. The indicative element 12 may be "fixed" below the parting line 11. For example, the indicative element 12 may be placed at a distance between one and five pixels away from the dividing line 11.

According to a seventh embodiment (fig. 11), the thickness E' of the dividing line 11 is equal to the pitch angle defined by the difference between the natural horizon and the horizon depending on the predetermined distance 17. The natural horizon corresponds to the horizon without the limitation of the field of view. The dividing line 11 includes pixels having values representing the third color range. The division line 11 is arranged at a pitch angle Δ' with respect to the zero pitch position of the aircraft AC, which is equal to the difference between zero pitch angle and the natural horizon.

In a first variant, the position of the horizon, which depends on the predetermined distance 17, is determined on the basis of the maximum distance defined by the maximum field of view of the topography of the terrain depicted in the second image when the aircraft AC is flying at maximum altitude. Therefore, the same equation as described in the fifth embodiment can be used to determine the position of the horizon, which is dependent on the predetermined distance 17, where the height is highDegree of rotation_FootCorresponding to the maximum altitude of the aircraft AC.

In a second variant, the position of the horizon, which depends on the predetermined distance 17, is determined based on the maximum distance defined by the maximum field of view of the topography of the terrain depicted in the second image, depending on the current altitude of the aircraft AC. Therefore, the same equation as described in the fifth embodiment can be used to determine the position of the horizon in which the height is high_footis variable.

the pitch angle Δ' depends on the altitude of the aircraft AC and may have the following function:

In this embodiment, the indicative element 12 may be "fixed" below the division line 11, for example, the indicative element 12 may be arranged at a distance between one and five pixels away from the division line 11. the above-described embodiment determines the pitch angles Δ ⁰ and Δ for a terrain flying with zero altitude (or in other words, the terrain surrounding the aircraft AC). The following embodiment determines the pitch angle Δ ⁰ while taking into account the altitude of the terrain surrounding the aircraft AC FIG. 7 shows the aircraft AC flying over the terrain, and an image I5. tending to be displayed in the display system of the aircraft AC. in this FIG. 7, the terrain surrounding the aircraft AC exhibits bumps 19, resulting in a non-zero altitude of the terrain surrounding the aircraft AC. in the case of FIG. 7, because the embodiment determining the position of the division line 11 assumes that the altitude of the terrain surrounding the aircraft AC is zero, the horizon line depending on the predetermined distance 17 rises relative to the division line 11.

The following embodiments provide a solution for addressing the curvature of the earth T and the limitations of the field of view 21 of the SVS while taking into account the altitude of the terrain surrounding the aircraft AC.

According to the eighth embodiment, the dividing line 11 is arranged at the pitch angle Δ with respect to the zero pitch position of the aircraft AC⁰where the pitch angle depends on the altitude of the aircraft 9 aircraft_FootMinus the altitude of the terrain over which the aircraft AC is flying 9 groundShape of_footand depends on the predetermined distance 17.

For example, the following equation allows obtaining the pitch angle Δ ⁰ from the altitude of the aircraft AC while taking into account the curvature of the earth T, the altitude at which the limits of the field of view 21 have been formed:

wherein:

-9 aircraft_Footis the current altitude (in feet) of the aircraft AC,

-9 topography_FootIs the altitude (in feet) of the terrain surrounding the aircraft AC, and

the SVS range _NM is a predetermined distance 17 (in nautical miles) defined by the maximum field of view 21,

-A, B and C are constants.

Equation (1) is applied under the following conditions:

Equation (2) is applied under the following conditions:

Without limitation, constant a equals 0.3048, constant B equals 1852, and constant C equals 1.23.

This embodiment is preferably implemented when the aircraft AC is close to the ground.

Thus, according to a variant embodiment, if the aircraft itself is in one of the following situations:

a final approach in the landing phase,

during the takeoff phase

-in the phase of the missed approach,

The altitude of the terrain flown over is equal to the current altitude of the terrain flown over in one of the aforementioned situations. For example, the altitude of the terrain is set at the altitude of the runway on which the aircraft AC lands, takes off, or performs a missed approach.

Thus, returning to the previous equation, the altitude Z terrain is determined by comparing the altitude Z terrain of the terrain surrounding the aircraft AC_FootDetermination of the pitch angle Δ of the dividing line 11 by replacement with the height of the runway⁰。

when the aircraft AC is not in one of the above-mentioned situations, the desensitizing means are implemented according to the previous embodiment in which it is assumed that the altitude of the terrain flown is zero.

The situation in which the aircraft AC itself is located may be determined from the altitude of the aircraft AC measured by the altimeter, from data representative of the flight phase of the aircraft AC provided by a system external to the apparatus, or from the distance between the current position of the aircraft AC and the departure or arrival point of the aircraft AC.

according to a variant of the eighth embodiment, the thickness E' of the dividing line 11 is equal to the pitch angle defined by the difference between the natural horizon and the horizon depending on the predetermined distance 17. The thickness E 'is determined in the same manner as the thickness E' of the seventh embodiment.

According to a ninth embodiment, the height Z topography of the terrain flown over_FootEqual to the minimum height of the terrain over which the aircraft AC flies within the section 21 of the cylinder 20 (fig. 13). The cylinder 20 is centered on the aircraft AC. The radius of the section 21 of the cylinder 20 is equal to the predetermined distance 17 and its aperture angle θ corresponds to the maximum field of view 21.

thus, returning to the previous equation, the terrain is determined by comparing the altitude 9 of the terrain surrounding the aircraft AC_FootThe pitch angle Δ of the division line 11 is determined by the minimum altitude of the terrain over which the aircraft AC flies within the section 21 of the cylinder 20⁰。

according to a variant of the ninth embodiment, the thickness E' of the dividing line 11 is equal to the pitch angle defined by the difference between the natural horizon and the horizon depending on the predetermined distance. The thickness E 'is determined in the same manner as the thickness E' of the seventh embodiment.

according to the tenth embodiment, the pitch angle Δ ⁰ of the division line 11 is determined directly from the apparent angle of the terrain visible from the position of the aircraft AC.

to this end, the section 21 of the cylinder 20 is divided into n discrete portions according to the angles θ ₁, …, θ _n (fig. 14).

For each discrete portion θ ₁, …, θ _n, the maximum terrain apparent angle α terrain _i is determined by the horizon hi order to obtain this angle α terrain _i for each discrete portion θ ₁, …, θ _i, …, θ _n, the height of the terrain may be determined at several points r ₁, r ₂, …, r _p, …, r _m along the discrete portion (fig. 15 and 16). the distance between the aircraft AC and point r _m corresponds to a predetermined distance 17. angle C terrain _i,p is then obtained.

Fig. 16 shows the angle C terrain _i,p determined for discrete portion θ _i and point r _p.

Thus, the angle C terrain _i,p is determined using the following equation:

for any discrete portion θ _i, angle C-feature _i is determined using the following equation:

the pitch Δ ⁰, which positions the dividing line 11, then corresponds to the minimum angle cbuture _i (fig. 17):

For the first embodiment, the indicative element 12 is arranged at a distance relative to the division line 11 representative of a pitch angle Δ ⁰ corresponding to the angle defined by the difference between the zero pitch angle of the aircraft AC and the natural horizon when the aircraft AC is flying at maximum altitude for other embodiments, the indicative element 12 may be "fixed" below the division line 11 according to the first configuration.

According to a second configuration (fig. 18 and 19), the indicative element 12 can be arranged at a distance corresponding to a pitch angle less than or equal to the pitch angle corresponding to the lowest point 23 of at least one ridge 24 seen in the display field of view 22, 27. Each ridge 24 corresponds to a natural horizon for a respective typical location on an approach trajectory, a missed approach trajectory, or a takeoff trajectory of the aircraft AC. For each of the various representative locations, each natural horizon corresponds to the horizon 18 in the image I2. The ridge line 24 is determined for a particular airport runway.

without limitation, the point potential is typically determined based on at least one of the following parameters:

-a selected approach runway (S),

-the current runway on which the aircraft AC may land,

-the altitude of the aircraft AC,

-the height of the selected runway threshold,

-lateral and vertical approach deviations,

-a vertical trajectory entry angle of field,

-a minimum decision altitude for the minimum decision altitude,

-maximum takeoff climb rate.

The display field of view 22 may correspond to the total field of view that tends to be displayed by the primary flight display PFD (fig. 18). For example, the display field of view is located between a speed scale 25(SPD) indicating the speed displayed in image I1 and an altitude scale 26(ALT) indicating the altitude displayed in image I1.

The display field of view 27 may also correspond to a portion of the total field of view intended to be displayed by the primary flight display PFD (fig. 19).

According to a first variant, the ridge 24 is predetermined and stored in a memory database of the aircraft AC. Advantageously, the form of the ridges 24 is discrete in order to minimize the memory required to store them.

According to a second embodiment, the ridges 24 are determined at each flight of the aircraft AC in order to be stored in a memory database of the aircraft AC or to update one or more ridges 24 already stored in the memory database.

The invention also relates to a method for desensitizing an aircraft AC display system configured to display a composite representation of terrain to faults.

The desensitization method (fig. 1) comprises the following steps:

an acquisition step E1, carried out by the acquisition module 3, comprising acquiring environmental data and current flight parameters of the aircraft AC, including at least attitude, altitude, three-dimensional position and heading;

-a generation step E2, carried out by the generation module 4, comprising generating an image I1 depicting a flight symbology comprising at least one zero-pitch reference line, the image I1 indicating a pitch angle with respect to a zero pitch angle of the aircraft AC, the zero-pitch reference line representing the zero pitch angle of the aircraft AC, the image I1 being generated at least on the basis of the attitude;

a generation step E3, carried out by the generation module 5, comprising generating an image I2, based on at least said attitude, said heading, said three-dimensional position and said environmental data, said image depicting the topography of the terrain over which the aircraft AC flies;

a generation step E4, carried out by generation module 6, comprising generating an image I3 comprising at least two parts separated by a dividing line determined on the basis of the attitude of aircraft AC, a first part 13 comprising pixels having values representative of a first color range, a second part 14 comprising pixels having values representative of a second color range;

a generation step E5, carried out by the generation module 7, comprising the generation of an image I4 by superimposing in sequence the image I1, the image I2 and the image I3, the image I1 being placed in the foreground;

a transmission step E6, carried out by the transmission module 8, comprising the transmission of a signal representative of the image I4 to the user device 9.

The generating step E3 may comprise the following sub-steps:

a determination sub-step E31, implemented by the determination sub-module 51, comprising determining the topography of the terrain over which the aircraft AC flies, based at least on said attitude, said three-dimensional position, said heading and said environmental data;

an assigning sub-step E32, implemented by the assigning sub-module 52, comprising assigning at least one pixel value to a first set of pixels forming part of the image I2, said one or more pixel values representing the topography of the terrain over which the aircraft AC flies;

A assigning sub-step E33, implemented by the assigning sub-module 53, comprising assigning values representative of transparent pixels to a second set of pixels forming part of the image I2, said second set of pixels corresponding to a set of pixels of the image I2 whose pixels do not belong to said first set of pixels.

The generating step E4 comprises the following sub-steps:

a determination sub-step E41, carried out by the determination sub-module 61, comprising determining the position of the dividing line, based on the zero-pitch reference line, the dividing line being parallel to the zero-pitch reference line 10, the portion 14 of the image I3 corresponding at least in part to the topography of the terrain 15 flown over by the aircraft AC represented in the image I2, when the image I1, the image I2 and the image I3 are superimposed in the generation step E5;

an adding sub-step E42, implemented by the adding module 62, comprising the addition of an indicative element 12 indicating an error to the portion 14 of the image I3.