CA2175030A1 - Aircraft instruments - Google Patents
Aircraft instrumentsInfo
- Publication number
- CA2175030A1 CA2175030A1 CA002175030A CA2175030A CA2175030A1 CA 2175030 A1 CA2175030 A1 CA 2175030A1 CA 002175030 A CA002175030 A CA 002175030A CA 2175030 A CA2175030 A CA 2175030A CA 2175030 A1 CA2175030 A1 CA 2175030A1
- Authority
- CA
- Canada
- Prior art keywords
- rate
- pitch angle
- aircraft
- change
- display
- Prior art date
- Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
- Abandoned
Links
- 230000002093 peripheral effect Effects 0.000 description 5
- 238000000034 method Methods 0.000 description 2
- 238000012935 Averaging Methods 0.000 description 1
- 230000001419 dependent effect Effects 0.000 description 1
- 238000010586 diagram Methods 0.000 description 1
- 230000000694 effects Effects 0.000 description 1
- 239000000446 fuel Substances 0.000 description 1
- 238000009434 installation Methods 0.000 description 1
- 239000004973 liquid crystal related substance Substances 0.000 description 1
- 239000011159 matrix material Substances 0.000 description 1
- 230000010355 oscillation Effects 0.000 description 1
- 239000000725 suspension Substances 0.000 description 1
- 230000000007 visual effect Effects 0.000 description 1
Classifications
-
- B—PERFORMING OPERATIONS; TRANSPORTING
- B64—AIRCRAFT; AVIATION; COSMONAUTICS
- B64D—EQUIPMENT FOR FITTING IN OR TO AIRCRAFT; FLIGHT SUITS; PARACHUTES; ARRANGEMENT OR MOUNTING OF POWER PLANTS OR PROPULSION TRANSMISSIONS IN AIRCRAFT
- B64D43/00—Arrangements or adaptations of instruments
-
- G—PHYSICS
- G05—CONTROLLING; REGULATING
- G05D—SYSTEMS FOR CONTROLLING OR REGULATING NON-ELECTRIC VARIABLES
- G05D1/00—Control of position, course, altitude or attitude of land, water, air or space vehicles, e.g. using automatic pilots
- G05D1/08—Control of attitude, i.e. control of roll, pitch, or yaw
- G05D1/0808—Control of attitude, i.e. control of roll, pitch, or yaw specially adapted for aircraft
- G05D1/0816—Control of attitude, i.e. control of roll, pitch, or yaw specially adapted for aircraft to ensure stability
Landscapes
- Engineering & Computer Science (AREA)
- Aviation & Aerospace Engineering (AREA)
- Radar, Positioning & Navigation (AREA)
- Remote Sensing (AREA)
- Physics & Mathematics (AREA)
- General Physics & Mathematics (AREA)
- Automation & Control Theory (AREA)
- Traffic Control Systems (AREA)
- Instrument Panels (AREA)
Abstract
The present invention relates to an aircraft instrument for reducing the risk of tail strike during aircraft take-off. The instrument receives an input from a pitch angle sensor and derives a signal indicative of rate of change of pitch angle. The instrument includes a comparator receiving the rate of change of pitch angle signal and signals from reference sources indicative of maximum and minimum rates of change of pitch angle. The instrument display is turned on when the aircraft nose gear lifts offthe runway and displays a stationary symbol as long as rate of change of pitch angle is within safe limits. When rate of change of pitch angle falls outside these limits, the display changes so that the symbol moves to warn the pilot. After the aircraft has reached a certain height, the instrument display is turned off.
Description
-Aircraft Instruments Back~round of the Invention This invention relates to aircraft instruments.
The invention is more particularly concerned with instruments for warning a pilot of an aircraft when the pitch rate deviates from a desired value during the rotation phase of take-off.
During take-offroll, the pilot rotates the aircraft (that is, lifts the nose offthe runway) when the aircraft's rotation speed, Vr, is reached. The climb attitude of an aircraft is typically 20 from the horizontal. Before the pilot can put the aircraft safely into this attitude, the aircraft must have lifted sufficiently to ensure that its tail does not strike the ground. Wing lift does not occur immediately that rotation takes place because of the time taken by air circulation to build up around the wing. The tail strike angle for a typical large wide-bodied airliner is about 10, or about 12 when the suspension of main wheel landing gear is extended as a result of the weight ofthe aircraft being taken offthe landing gear during rotation. With an extended version of the aircraft, these angles would be reduced to about 7 and 8 respectively. Tail strike can occur on the ground if the nose is raised too quickly before the main landing gear has lifted offthe ground. Tail strike can also occur in the air if the nose is raised too quickly before sufficient height has been reached. The pilot must, therefore, ensure that he does not rotate the aircraft with too high a pitch rate, to avoid tail strike.
Although instruments have been proposed previously for providing information to the pilot during the flare-up phase oftake-off, such as described in US3309923, no instrument has previously been able to provide information warning of possible tail strike during the very early phase of take-off.
The take-off procedure is particularly stressful for the pilot because the engines are operating at full power and the plane is fully loaded with fuel. The pilot must monitor many instruments as well as controlling the engines and aerodynamic surfaces of the aircraft.
Brief Summary of the Present Invention It is an object of the present invention to provide an aircraft instrument that can be used to assist the pilot during the rotation phase of take-off.
According to the present invention there is provided an aircraft instrument including means for providing a signal indicative of the rate of change of pitch angle of the aircraft, means for determining when the rate of change of pitch angle departs from a safe value, and display means for providing a warning display to the pilot when the rate of change of pitch angle during the rotation phase of take-off departs from the safe value.
The instrument preferably includes means for receiving an input from a sensor responsive to lifting of the aircraft nose wheel at the start of rotation. The instrument is preferably arranged to provide a first warning display when the pitch rate is too high and a di~el ellL warning display when the pitch rate is too low. The warning display may include a representation of a symbol that moves vertically when the pitch rate departs from a safe value and that remains stationary when the pitch rate is at a safe value. The instrument may include a processor, comparator means, a reference source of maximum rate of change of pitch angle and a reference source of miniml]m rate of change of pitch angle, the processor being arranged to receive signals from a pitch angle sensor and to derive a signal indicative of rate of change of pitch angle, the comparator means receiving outputs from the reference sources of minimllm and maximum rate of change of pitch angle and the signal indicative of rate of change of pitch angle, and the comparator means providing an output to initiate said warning display when the signal indicative of rate of change of pitch angle exceeds the maximum rate of change of pitch angle or falls below the minimllm rate of change of pitch angle. The instrument may be arranged to m~int~in the display offuntil the aircraft nose wheel lifts offthe ground. The instrument may be arranged to turn offthe display a predetermined time after the main landing gear has lifted offthe ground or after the aircraft has reached a predetermined height above the ground. The instrument may be arranged to provide an audible warning when the rate of change of pitch angle during the rotation phase of take-off departs from a safe value. The display of the instrument is preferably arranged for mounting in the peripheral field of view of the pilot. The instrument may also be arranged to provide a warning display during descent if pitch angle of the aircraft exceeds a safe value, and to provide lateral guidance information to the pilot.
A tail-strike warning instrument for an aircraft, in accordance with the presentinvention, will now be described, by way of example, with reference to the accompanying drawings.
Brief Description of the Drawin~s Figure 1 is a schematic illustration of an aircraft showing the installation of the instrument and various sensors;
Figure 2 is a schematic diagram of the instrument;
Figures 3A to 3 C show di~eren~ display representations provided by the instrument; and Figures 4 to 7 show alternative formats of display representation.
The invention is more particularly concerned with instruments for warning a pilot of an aircraft when the pitch rate deviates from a desired value during the rotation phase of take-off.
During take-offroll, the pilot rotates the aircraft (that is, lifts the nose offthe runway) when the aircraft's rotation speed, Vr, is reached. The climb attitude of an aircraft is typically 20 from the horizontal. Before the pilot can put the aircraft safely into this attitude, the aircraft must have lifted sufficiently to ensure that its tail does not strike the ground. Wing lift does not occur immediately that rotation takes place because of the time taken by air circulation to build up around the wing. The tail strike angle for a typical large wide-bodied airliner is about 10, or about 12 when the suspension of main wheel landing gear is extended as a result of the weight ofthe aircraft being taken offthe landing gear during rotation. With an extended version of the aircraft, these angles would be reduced to about 7 and 8 respectively. Tail strike can occur on the ground if the nose is raised too quickly before the main landing gear has lifted offthe ground. Tail strike can also occur in the air if the nose is raised too quickly before sufficient height has been reached. The pilot must, therefore, ensure that he does not rotate the aircraft with too high a pitch rate, to avoid tail strike.
Although instruments have been proposed previously for providing information to the pilot during the flare-up phase oftake-off, such as described in US3309923, no instrument has previously been able to provide information warning of possible tail strike during the very early phase of take-off.
The take-off procedure is particularly stressful for the pilot because the engines are operating at full power and the plane is fully loaded with fuel. The pilot must monitor many instruments as well as controlling the engines and aerodynamic surfaces of the aircraft.
Brief Summary of the Present Invention It is an object of the present invention to provide an aircraft instrument that can be used to assist the pilot during the rotation phase of take-off.
According to the present invention there is provided an aircraft instrument including means for providing a signal indicative of the rate of change of pitch angle of the aircraft, means for determining when the rate of change of pitch angle departs from a safe value, and display means for providing a warning display to the pilot when the rate of change of pitch angle during the rotation phase of take-off departs from the safe value.
The instrument preferably includes means for receiving an input from a sensor responsive to lifting of the aircraft nose wheel at the start of rotation. The instrument is preferably arranged to provide a first warning display when the pitch rate is too high and a di~el ellL warning display when the pitch rate is too low. The warning display may include a representation of a symbol that moves vertically when the pitch rate departs from a safe value and that remains stationary when the pitch rate is at a safe value. The instrument may include a processor, comparator means, a reference source of maximum rate of change of pitch angle and a reference source of miniml]m rate of change of pitch angle, the processor being arranged to receive signals from a pitch angle sensor and to derive a signal indicative of rate of change of pitch angle, the comparator means receiving outputs from the reference sources of minimllm and maximum rate of change of pitch angle and the signal indicative of rate of change of pitch angle, and the comparator means providing an output to initiate said warning display when the signal indicative of rate of change of pitch angle exceeds the maximum rate of change of pitch angle or falls below the minimllm rate of change of pitch angle. The instrument may be arranged to m~int~in the display offuntil the aircraft nose wheel lifts offthe ground. The instrument may be arranged to turn offthe display a predetermined time after the main landing gear has lifted offthe ground or after the aircraft has reached a predetermined height above the ground. The instrument may be arranged to provide an audible warning when the rate of change of pitch angle during the rotation phase of take-off departs from a safe value. The display of the instrument is preferably arranged for mounting in the peripheral field of view of the pilot. The instrument may also be arranged to provide a warning display during descent if pitch angle of the aircraft exceeds a safe value, and to provide lateral guidance information to the pilot.
A tail-strike warning instrument for an aircraft, in accordance with the presentinvention, will now be described, by way of example, with reference to the accompanying drawings.
Brief Description of the Drawin~s Figure 1 is a schematic illustration of an aircraft showing the installation of the instrument and various sensors;
Figure 2 is a schematic diagram of the instrument;
Figures 3A to 3 C show di~eren~ display representations provided by the instrument; and Figures 4 to 7 show alternative formats of display representation.
Detailed Description of the Preferred Embodiments With reference first to Figures 1 and 2, the instrument includes an electronics housing 1 connected to a display unit 2 by a cable 3. The display unit 2 is mounted in the glareshield of the aircraft flight deck so that it is in the peripheral field of view of the pilot; the housing 1 may be mounted anywhere in the aircraft or within the same unit as the display itself. The instrument receives inputs from a nosegear squat switch 4 and from a main gear squat switch 5. These switches 4 and 5 provide outputs to indicate whether or not the nose gear or main gear are on the ground. The instrument also receives an output ~ from a pitch angle sensor 6.
This sensor 6 may be contained within the electronics housing 1 or it could be a discrete sensor located externally. Alternatively, the pitch angle sensing function could be provided by an existing pitch angle sensor used for other purposes.
The electronics housing 1 includes a processor 40, which receives the output from the pitch angle sensor 6 and, from this, derives an output ~ indicative of the rate of change of pitch angle. Alternatively, pitch rate may be input directly from a pitch rate gyroscope sensor.
The processor 40 preferably also performs an averaging function to reduce the effect of pilot-induced oscillations or other small perturbations in the pitch angle signal. The output from the processor 40 is connected to one input of each of two comparators 41 and 42. Onecomparator 41 has its other input connected to a reference source 43, which sets a maximum value of rate of change of pitch angle ~ . The other comparator 42 has its other input connected to a reference source 44, which sets a minimllm value of rate of change of pitch angle ~ ,. The outputs from the two comparators 41 and 42 are connected to a display driver unit 50. The display driver unit 50 also receives the outputs from the nosegear squat switch 4 and from the main gear squat switch 5. The display driver unit 50 is connected to the display unit 2 by the cable 3 and provides the output from the electronics housing 1.
The display unit 2 is of rectangular shape and has a front surface or screen 20 occupied by a matrix array of liquid crystal display elements 21, or some other electrically-energizable display elements, such as LEDs.
When the aircraft starts its ground roll, both its nose and main gear are on the ground and the sensors 4 and 5 supply signals indicating this to the instrument 1. During this part of the take-off procedure, the instrument 1 holds the display unit 2 off so that the pilot is not distracted.
When the pilot pulls back on the stick to raise the nose of the aircraft and start the rotation phase of take-off, the nose gear starts to lift away from the ground and the nosegear squat switch 4 changes its output. This causes the instrument 1 to energize the display unit 2.
While the pilot m~in~in.~ the pitch rate of the aircraft within safe limits, the display driver 50 produces a display representation on the screen 20 of the kind shown in Figure 3B. This comprises a number of dark horizontal bars 22 (three bars are shown in Figure 3B) extending across the display and separated by bright gaps 23. This display representation remains stationary while the aircraft is m~int~ined within safe pitch rate lirnits, that is, less than ~3maX and more than ~"il,. The pilot will see the display as it is turned on, in his peripheral field of view, so that he is notified that the nose gear has lifted offthe runway. During rotation, the pilot receives guidance from the display in his peripheral field of vision while looking forward out of the cockpit windscreen, and without having to focus his eyes on the display.
When the main gear of the aircraft lifts offthe runway, the sensor S changes its output.
This causes the processor 40 to start a timer and, after a predetermined time has elapsed suffficient for the aircraft to have achieved a height at which tail strike is no danger (typically about S seconds) the instrument 1 turns offthe display 2, which is no longer needed.
Alternatively, the instrument could be connected to receive an output from the aircraft's radar altimeter 7, instead of from a main gear squat switch 5. In this case, the display unit 2 would be turned offwhen the aircraft has achieved a safe height at which there is no risk of tail strike.
If, however, the pilot were to pull back on the stick too quickly and cause an excessively high pitch rate, sufficient for there to be a danger of tail strike, the input to the comparator 41 would exceed the reference value ~ma~ from the source 43. The comparator 41 would then change its output to the display driver unit 50 so that the driver unit displaces the bars 22 downwards, giving an appearance of a continuous stream of bars moving down. The rate of movement of the bars is proportional to the magnitude of the difference between the actual aircraft pitch rate and the maximum safe pitch rate. This warning movement on the display is readily apparent to the pilot in his peripheral field of view without him having to look directly at the display. The pilot can, therefore, immediately take correcting action without having to turn his head or alter his focus. The pilot will notice that his correcting action produces a slowing down of the moving bars until the aircraft comes below the upper safe pitch rate limit, when the display representation again becomes stationary.
If the pilot were over cautious and did not produce a sufficiently high pitch rate, there would be a risk that the aircraft would not produce sufficient lift quickly enough and might run out of runway. When the pitch rate is too low, the input the processor 40 supplies to the comparator 42 falls below that from the reference source 43. This causes the comparator 42 to change its output, which, in turn, causes the display driver unit 50 to move the bars on the display representation upwardly at a rate proportional to the m~Enit~lde of the difference between the actual aircraft pitch rate and the minimum safe pitch rate ~mi". This warns the pilot that he must increase the pitch rate.
The present invention enables the pilot to be warned when the pitch rate of the aircraft is outside safe lirnits, without him being presented with distracting information when it is not needed.
The instrument could also have an audible warning, which it triggers when the visual warning does not produces a corrective response by the pilot within a predetermined time.
The display representation could take various other forms such as, for example, shown in Figures 4 to 7. In Figure 4B, the safe pitch rate is represented by a continuous, horizontal line extending across the display midway up its height. When the pitch rate is too high, the central part of the line is displaced down, as shown in Figure 4A; when the pitch rate is too low, the central part of the line is displaced up, as shown in Figure 4C. In the display representation shown in Figure 5, the correct pitch rate produces a blank display, as shown in Figure 5B, whereas too fast a pitch rate produces a downwardly-pointing arrow, as shown in Figure 5A, and too slow a pitch rate produces the upwardly-directed arrow, as shown in Figure 5C. The display representation could be arranged to change colour, as shown in Figure 6. When the pitch rate is within a safe range, the display representation is a plain green screen, as shown in Figure 6B; when the pitch rate is too high, the display changes to a red colour and text, such as "FAST" appears on the display, as shown in Figure 6A; and when the pitch rate is too low, the display changes to an amber colour and text, such as the word "SLOW" is displayed. The display format of Figure 6 could be combined with a moving representation so that the pilot's attention is drawn to the display more forcefully.
The instrument could also be used to provide a warning of tail strike during landing as a result of too large a pitch angle during descent. The risk of tail strike during landing depends only on pitch attitude, rather than pitch rate, so the processor does not compute the rate of change of pitch angle during this phase.
In the display shown in Figure 7, the upper part of the display screen 70 shows a display representation of the same kind as that in Figure 6. The lower part of the screen 71 is occupied by a lateral guidance display. The lateral guidance display indicates to the pilot if the aircraft deviates from the runway centre line. The lateral guidance display is formed by inclined stripes, which remain stationary when the aircraft is correctly aligned with the runway.
When the aircraft heading deviates to right or left of the centre line, the stripes move across the width of the display to the right or left accordingly and at a rate dependent on the magnitude of the deviation.
Because the tail strike warning instrument is only used during the rotation phase of take-off, or during take-off and landing, it is possible for the display to be used for other purposes at other times. For example, it could be used to display a warning of collision avoidance action to be taken when there is a risk of a mid-air collision, in the manner described in GB 2226924. Alternatively, it could be used to display air traffic command instructions, as described in GB 2250494.
This sensor 6 may be contained within the electronics housing 1 or it could be a discrete sensor located externally. Alternatively, the pitch angle sensing function could be provided by an existing pitch angle sensor used for other purposes.
The electronics housing 1 includes a processor 40, which receives the output from the pitch angle sensor 6 and, from this, derives an output ~ indicative of the rate of change of pitch angle. Alternatively, pitch rate may be input directly from a pitch rate gyroscope sensor.
The processor 40 preferably also performs an averaging function to reduce the effect of pilot-induced oscillations or other small perturbations in the pitch angle signal. The output from the processor 40 is connected to one input of each of two comparators 41 and 42. Onecomparator 41 has its other input connected to a reference source 43, which sets a maximum value of rate of change of pitch angle ~ . The other comparator 42 has its other input connected to a reference source 44, which sets a minimllm value of rate of change of pitch angle ~ ,. The outputs from the two comparators 41 and 42 are connected to a display driver unit 50. The display driver unit 50 also receives the outputs from the nosegear squat switch 4 and from the main gear squat switch 5. The display driver unit 50 is connected to the display unit 2 by the cable 3 and provides the output from the electronics housing 1.
The display unit 2 is of rectangular shape and has a front surface or screen 20 occupied by a matrix array of liquid crystal display elements 21, or some other electrically-energizable display elements, such as LEDs.
When the aircraft starts its ground roll, both its nose and main gear are on the ground and the sensors 4 and 5 supply signals indicating this to the instrument 1. During this part of the take-off procedure, the instrument 1 holds the display unit 2 off so that the pilot is not distracted.
When the pilot pulls back on the stick to raise the nose of the aircraft and start the rotation phase of take-off, the nose gear starts to lift away from the ground and the nosegear squat switch 4 changes its output. This causes the instrument 1 to energize the display unit 2.
While the pilot m~in~in.~ the pitch rate of the aircraft within safe limits, the display driver 50 produces a display representation on the screen 20 of the kind shown in Figure 3B. This comprises a number of dark horizontal bars 22 (three bars are shown in Figure 3B) extending across the display and separated by bright gaps 23. This display representation remains stationary while the aircraft is m~int~ined within safe pitch rate lirnits, that is, less than ~3maX and more than ~"il,. The pilot will see the display as it is turned on, in his peripheral field of view, so that he is notified that the nose gear has lifted offthe runway. During rotation, the pilot receives guidance from the display in his peripheral field of vision while looking forward out of the cockpit windscreen, and without having to focus his eyes on the display.
When the main gear of the aircraft lifts offthe runway, the sensor S changes its output.
This causes the processor 40 to start a timer and, after a predetermined time has elapsed suffficient for the aircraft to have achieved a height at which tail strike is no danger (typically about S seconds) the instrument 1 turns offthe display 2, which is no longer needed.
Alternatively, the instrument could be connected to receive an output from the aircraft's radar altimeter 7, instead of from a main gear squat switch 5. In this case, the display unit 2 would be turned offwhen the aircraft has achieved a safe height at which there is no risk of tail strike.
If, however, the pilot were to pull back on the stick too quickly and cause an excessively high pitch rate, sufficient for there to be a danger of tail strike, the input to the comparator 41 would exceed the reference value ~ma~ from the source 43. The comparator 41 would then change its output to the display driver unit 50 so that the driver unit displaces the bars 22 downwards, giving an appearance of a continuous stream of bars moving down. The rate of movement of the bars is proportional to the magnitude of the difference between the actual aircraft pitch rate and the maximum safe pitch rate. This warning movement on the display is readily apparent to the pilot in his peripheral field of view without him having to look directly at the display. The pilot can, therefore, immediately take correcting action without having to turn his head or alter his focus. The pilot will notice that his correcting action produces a slowing down of the moving bars until the aircraft comes below the upper safe pitch rate limit, when the display representation again becomes stationary.
If the pilot were over cautious and did not produce a sufficiently high pitch rate, there would be a risk that the aircraft would not produce sufficient lift quickly enough and might run out of runway. When the pitch rate is too low, the input the processor 40 supplies to the comparator 42 falls below that from the reference source 43. This causes the comparator 42 to change its output, which, in turn, causes the display driver unit 50 to move the bars on the display representation upwardly at a rate proportional to the m~Enit~lde of the difference between the actual aircraft pitch rate and the minimum safe pitch rate ~mi". This warns the pilot that he must increase the pitch rate.
The present invention enables the pilot to be warned when the pitch rate of the aircraft is outside safe lirnits, without him being presented with distracting information when it is not needed.
The instrument could also have an audible warning, which it triggers when the visual warning does not produces a corrective response by the pilot within a predetermined time.
The display representation could take various other forms such as, for example, shown in Figures 4 to 7. In Figure 4B, the safe pitch rate is represented by a continuous, horizontal line extending across the display midway up its height. When the pitch rate is too high, the central part of the line is displaced down, as shown in Figure 4A; when the pitch rate is too low, the central part of the line is displaced up, as shown in Figure 4C. In the display representation shown in Figure 5, the correct pitch rate produces a blank display, as shown in Figure 5B, whereas too fast a pitch rate produces a downwardly-pointing arrow, as shown in Figure 5A, and too slow a pitch rate produces the upwardly-directed arrow, as shown in Figure 5C. The display representation could be arranged to change colour, as shown in Figure 6. When the pitch rate is within a safe range, the display representation is a plain green screen, as shown in Figure 6B; when the pitch rate is too high, the display changes to a red colour and text, such as "FAST" appears on the display, as shown in Figure 6A; and when the pitch rate is too low, the display changes to an amber colour and text, such as the word "SLOW" is displayed. The display format of Figure 6 could be combined with a moving representation so that the pilot's attention is drawn to the display more forcefully.
The instrument could also be used to provide a warning of tail strike during landing as a result of too large a pitch angle during descent. The risk of tail strike during landing depends only on pitch attitude, rather than pitch rate, so the processor does not compute the rate of change of pitch angle during this phase.
In the display shown in Figure 7, the upper part of the display screen 70 shows a display representation of the same kind as that in Figure 6. The lower part of the screen 71 is occupied by a lateral guidance display. The lateral guidance display indicates to the pilot if the aircraft deviates from the runway centre line. The lateral guidance display is formed by inclined stripes, which remain stationary when the aircraft is correctly aligned with the runway.
When the aircraft heading deviates to right or left of the centre line, the stripes move across the width of the display to the right or left accordingly and at a rate dependent on the magnitude of the deviation.
Because the tail strike warning instrument is only used during the rotation phase of take-off, or during take-off and landing, it is possible for the display to be used for other purposes at other times. For example, it could be used to display a warning of collision avoidance action to be taken when there is a risk of a mid-air collision, in the manner described in GB 2226924. Alternatively, it could be used to display air traffic command instructions, as described in GB 2250494.
Claims (12)
1. An aircraft instrument comprising: a first unit, said first unit receiving an input indicative of aircraft pitch angle and providing an output indicative of the rate of change of pitch angle; a second unit connected with said first unit, said second unit determining when the rate of change of pitch angle during the rotation phase of take-off departs from a safe value; and a display connected to said second unit, said display providing a warning display to a pilot when the rate of change of pitch angle during the rotation phase of take-off departs from said safe value.
2. An aircraft instrument according to Claim 1, wherein said instrument has an input connected to a sensor responsive to lifting of a nose wheel of the aircraft at the start of rotation.
3. An aircraft instrument according to Claim 1 or 2, wherein said instrument provides a first warning display when the pitch rate is too high and a different warning display when the pitch rate is too low.
4. An aircraft instrument according to Claim 1 or 2, wherein the warning display includes a representation of a symbol that moves vertically when the pitch rate departs from a safe value and that remains stationary when the pitch rate is at a safe value.
5. An aircraft instrument according to Claim 1, wherein the instrument maintains the display off until the aircraft nose wheel lifts off the ground.
6. An aircraft instrument according to Claim 1, wherein said instrument is arranged to turn off said display a predetermined time after the main landing gear has lifted off the ground.
7. An aircraft instrument according to Claim 1, wherein said instrument is arranged to turn off said display after the aircraft has reached a predetermined height above the ground.
8. An aircraft instrument according to Claim 1, wherein said instrument provides an audible warning when the rate of change of pitch angle during the rotation phase of take-off departs from a safe value.
9. An aircraft instrument according to Claim 1, wherein said instrument is also arranged to provide a warning display during descent if the pitch angle of the aircraft exceeds a safe value.
10. An aircraft instrument according to Claim 1, wherein said instrument is also arranged to provide lateral guidance information to the pilot.
11. An aircraft instrument comprising: a display; a processor, said processor having an input connected to a pitch angle sensor and said processor providing an output signal indicative of rate of change of pitch angle; a comparator; a first reference source of a maximum rate of change of pitch angle; a second reference source of a minimum rate of change of pitch angle; a connection between an input of said comparator and an output of said processor; and a connection between outputs from said first and second reference sources and an input of said comparator, said comparator determining when the rate of change of pitch angle during the rotation phase of take-off is more than said maximum or less than said minimum rate of change of pitch angle and providing an output accordingly to initiate a warning on said display.
12. An aircraft instrument comprising: a display; means for providing an output signal indicative of rate of change of pitch angle; a comparator; a first reference source of a maximum rate of change of pitch angle; a second reference source of a minimum rate of change of pitch angle; a connection between an input of said comparator and said means for providing an output signal indicative of rate of change of pitch angle; and a connection between outputs from said first and second reference sources and an input of said comparator, said comparator determining when the rate of change of pitch angle during the rotation phase of take-off is more than said maximum or less than said minimum rate of change of pitch angle and providing an output accordingly to initiate a warning on said display.
Applications Claiming Priority (2)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| GBGB9508659.1A GB9508659D0 (en) | 1995-04-28 | 1995-04-28 | Aircraft instruments |
| GB9508659 | 1995-04-28 |
Publications (1)
| Publication Number | Publication Date |
|---|---|
| CA2175030A1 true CA2175030A1 (en) | 1996-10-29 |
Family
ID=10773680
Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| CA002175030A Abandoned CA2175030A1 (en) | 1995-04-28 | 1996-04-25 | Aircraft instruments |
Country Status (4)
| Country | Link |
|---|---|
| CA (1) | CA2175030A1 (en) |
| DE (1) | DE19615258A1 (en) |
| FR (1) | FR2733597B1 (en) |
| GB (2) | GB9508659D0 (en) |
Cited By (1)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN113286087A (en) * | 2021-05-28 | 2021-08-20 | 杭州微影软件有限公司 | Screen control method and device and thermal imager |
Families Citing this family (7)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| JP2939234B1 (en) | 1998-03-24 | 1999-08-25 | 株式会社コミュータヘリコプタ先進技術研究所 | Flight path display device |
| DE19930559B4 (en) * | 1999-07-02 | 2004-08-12 | Airbus Deutschland Gmbh | Arrangement and method for protecting an aircraft fuselage |
| GB2510608B (en) * | 2013-02-08 | 2015-02-25 | Ge Aviat Systems Ltd | Method for predicting a horizontal stabilizer fault |
| US9828113B2 (en) | 2013-11-05 | 2017-11-28 | Safe Flight Instrument Corporation | Tailstrike warning system |
| EP3066651A4 (en) * | 2013-11-05 | 2017-07-19 | Safe Flight Instrument Corporation | Tailstrike warning system |
| US20170008639A1 (en) | 2015-07-08 | 2017-01-12 | Safe Flight Instrument Corporation | Aircraft turbulence detection |
| CN110069070B (en) * | 2019-05-08 | 2022-01-18 | 成都高威节能科技有限公司 | Method for improving safety of large airplane in takeoff process |
Family Cites Families (12)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| GB1032466A (en) * | 1962-11-20 | 1966-06-08 | Smiths Industries Ltd | Improvements in or relating to aircraft instruments |
| GB1042170A (en) * | 1963-11-28 | 1966-09-14 | Smiths Industries Ltd | Improvements in or relating to aircraft instruments |
| SE335248B (en) * | 1970-05-11 | 1971-05-17 | Saab Scania Ab | |
| US4043526A (en) * | 1976-02-23 | 1977-08-23 | The United States Of America As Represented By The Secretary Of The Navy | Autopilot hardover failure protection system |
| US4046993A (en) * | 1976-06-28 | 1977-09-06 | The United States Of America As Represented By The Secretary Of The Navy | Target for torpedo launch system |
| US4071893A (en) * | 1976-07-06 | 1978-01-31 | Societe Francaise D'equipements Pour La Navigation Aerienne | Flying method and system using total power for an aircraft |
| GB2134866B (en) * | 1980-11-28 | 1985-06-19 | Sundstrand Data Control | Angle of attack based pitch generator and head up display |
| GB2179612B (en) * | 1982-07-12 | 1987-08-26 | Secr Defence | Aircraft instrumentation |
| US4769645A (en) * | 1983-06-10 | 1988-09-06 | Sundstrand Data Control, Inc. | Excessive pitch attitude warning system for rotary wing aircraft |
| EP0224278B1 (en) * | 1985-11-20 | 1991-05-08 | The Boeing Company | Apparatus for generating an aircraft situation display |
| GB9003653D0 (en) * | 1990-02-17 | 1990-04-11 | Smiths Industries Plc | Aircraft performance monitoring |
| US5169090A (en) * | 1991-08-28 | 1992-12-08 | United Technologies Corporation | Attitude synchronization for model following control systems |
-
1995
- 1995-04-28 GB GBGB9508659.1A patent/GB9508659D0/en active Pending
-
1996
- 1996-04-10 GB GB9607429A patent/GB2300167B/en not_active Expired - Fee Related
- 1996-04-18 DE DE19615258A patent/DE19615258A1/en not_active Withdrawn
- 1996-04-25 CA CA002175030A patent/CA2175030A1/en not_active Abandoned
- 1996-04-25 FR FR9605587A patent/FR2733597B1/en not_active Expired - Fee Related
Cited By (2)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN113286087A (en) * | 2021-05-28 | 2021-08-20 | 杭州微影软件有限公司 | Screen control method and device and thermal imager |
| CN113286087B (en) * | 2021-05-28 | 2022-09-02 | 杭州微影软件有限公司 | Screen control method and device and thermal imager |
Also Published As
| Publication number | Publication date |
|---|---|
| GB9508659D0 (en) | 1995-06-14 |
| DE19615258A1 (en) | 1996-10-31 |
| GB2300167B (en) | 1999-10-06 |
| FR2733597B1 (en) | 1999-08-13 |
| GB9607429D0 (en) | 1996-06-12 |
| FR2733597A1 (en) | 1996-10-31 |
| GB2300167A (en) | 1996-10-30 |
Similar Documents
| Publication | Publication Date | Title |
|---|---|---|
| US6702229B2 (en) | Method, apparatus and article to display flight information | |
| US6107943A (en) | Display symbology indicating aircraft ground motion deceleration | |
| US8421649B2 (en) | Aircraft attitude systems | |
| US4484191A (en) | Tactile signaling systems for aircraft | |
| US20130274965A1 (en) | Automated take off control system and method | |
| US11365971B2 (en) | Aircraft energy state awareness display systems and methods | |
| US8290641B2 (en) | Aircraft attitude systems and related methods | |
| JP5185141B2 (en) | Method and apparatus for automatically adjusting aircraft navigation screen images | |
| JPS63503093A (en) | Wind shear detection head-up display method | |
| US20150084792A1 (en) | Angle of attack display | |
| EP0817952B1 (en) | Aircraft flight instrument displays | |
| JP2013237434A (en) | Aircraft and method for displaying visual information associated to flight parameter to operator of aircraft | |
| US20190004081A1 (en) | Sideslip guidance for one engine inoperative condition | |
| CA2175030A1 (en) | Aircraft instruments | |
| US5675328A (en) | Optoelectronic device for assistance in the piloting of an aircraft under conditions of poor visibility | |
| US8224506B2 (en) | Method and device for determining a maximum stabilization height in the final flight phase of an airplane | |
| US20180022469A1 (en) | Head-Up Display (HUD) Stall Recovery Symbology | |
| EP0708394B1 (en) | Optoelectronic device for assisting a pilot in steering an aircraft | |
| EP0224278B1 (en) | Apparatus for generating an aircraft situation display | |
| RU2241642C2 (en) | Method and device for piloting of aircraft and aircraft | |
| GB2139588A (en) | System for alerting a pilot of a dangerous flight profile during low level maneuvering | |
| RU2653414C1 (en) | Stalling warning system | |
| US8774986B1 (en) | Method, system, and apparatus for takeoff rotation guidance | |
| RU2729891C1 (en) | Intelligent man-machine interface of helicopter crew on altitude-speed parameters and parameters of air environment surrounding helicopter | |
| US9218743B2 (en) | Navigation aid instrument for aircraft |
Legal Events
| Date | Code | Title | Description |
|---|---|---|---|
| FZDE | Discontinued |