Title of Invention

AIRCRAFT IDENTIFICATION AND DOCKING GUIDANCE SYSTEMS

Abstract A laser range finder (LRF) (20) is used to identify an aircraft (12) approaching a gate. The LRF (20) is directed at the aircraft (12), and from the echoes, a profile is derived and compared to known profiles. To distinguish among aircraft (12) with similar profiles, the LRF is directed at a volume in which a feature such as an engine is expected and at another volume in which the engine is not expected. The echoes from the two volumes are used to determine whether the engine is in its expected location. If so, the aircraft (12) is identified as the correct type and is allowed to dock at the gate (16). Otherwise, the aircraft (12) is stopped. The nose height can be used as yet another identifying criterion.
Full Text AIRCRAFT IDENTIFICATION AND
DOCKING GUIDANCE SYSTEMS
REFERENCE TO RELATED APPLICATIONS
This is a continuation-in-part of U.S. Patent Application No. 09/429,609, filed October
29, 1999, currently pending, which is a continuation-in-part of U.S. Patent Application No.
08/817,368, filed July 17, 1997, now U.S. Patent No. 6,023,665, which is the U.S. national stage
of PCT International Application No. PCT/SE94/00968, filed October 14,1994, published April
25, 1996, as WO 96/12265 Al. The disclosures of the parent applications are hereby
incorporated by reference in their entireties into the present disclosure.
BACKGROUND OF THE INVENTION
Field of the Invention
This invention relates to systems for locating, identifying and tracking objects. More
particularly, it related to aircraft location, identification and docking guidance systems and to
ground traffic control methods for locating and identifying objects on an airfield and for safely
and efficiently docking aircraft at such airport.
Description of Related Art
In recent years there has been a significantly increased amount of passenger, cargo and
other aircraft traffic including take offs, landings and other aircraft ground traffic. Also, there
has been a marked increase in the number of ground support vehicles which are required to off
load cargo, provide catering services and on going maintenance and support of all aircraft. With
that substantial increase in ground traffic has come a need for greater control and safety in the
docking and identification, of aircraft on an airfield.
Exemplary of prior art systems which have been proposed for detecting the presence of
aircraft and other traffic on an airfield are those systems disclosed in U.S. Patent 4,995,102;
European Patent No. 188 757: and PCT Published Applications WO 93/13104 and WO
93/15416.
However, none of those systems have been found to be satisfactory for detection of the
presence of aircraft on an airfield, particularly, under adverse climatic conditions causing
diminished visibility such as encountered under fog, snow or sleet conditions. Furthermore, none
of the systems disclosed in the prior references are capable of identifying and verifying the
specific type of an approaching aircraft. Still further, none of the prior systems provide adequate
techniques for tracking and docking an aircraft at a designated stopping point such as an airport
loading gate. Also, none of the prior systems have provided techniques which enable adequate
calibration of the instrument therein.
The system disclosed in the above-cited parent application seeks to overcome the above-
noted problems though profile matching. Light pulses from a laser range finder (LRF) are
projected in angular coordinates onto the airplane. The light pulses are reflected off the airplane
to detect a shape of the airplane or of a portion of the airplane, e.g., the nose. The detected shape
is compared with a profile corresponding to the shape of a known model of airplane to determine
whether the detected shape corresponds to the shape of the known model.
However, that system has a drawback. Often, two or more models of airplanes have nose
profiles so similar that one model is often misidentified as another. In particular, in adverse
weather, many echoes are lost, so that profile discrimination becomes decreasingly reliable.
Since the models are similar but not identical in body configuration, a correct docking position
for one can cause an engine on another to crash into a physical obstacle.
Thus, it has been a continuing problem to provide systems which are sufficiently safe and
reliable over a wide range of atmospheric conditions to enable detection of objects such as
aircraft and other ground traffic on an airfield.
In addition, there has been a long standing need for systems which are not only capable
of detecting objects such as aircraft, but which also provide for the effective identification of the
detected object and verification of the identity of such object, for example, a detected aircraft
with the necessary degree of certainty regardless of prevailing weather conditions and magnitude
of ground traffic.
There has also been a long standing, unfulfilled need for systems which are capable of
accurately and efficiently tracking and guiding objects such as incoming aircraft to a suitable
stopping point such as an airport loading gate. In addition, the provision of accurate and
effective calibration techniques for such systems has been a continuing problem requiring
resolution.
SUMMARY OF THE INVENTION
It will be readily apparent from the above that a need exists in the art for a more accurate
identification of aircraft.
It is therefore a primary object of the invention to distinguish among multiple models of
aircraft with identical or almost identical nose shapes.
It is a further object of the invention to improve the detection of aircraft so as to avoid
accidents during aircraft docking.
To achieve the above and other objects, the present invention identifies aircraft in a two-
step process. First, the profile matching is performed as known from the above-identified parent
application. Second, at least one aircraft criterion matching is performed. In the aircraft criterion
matching, a component of the aircraft, such as the engine, is selected as a basis for distinguishing
among aircraft. The displacement of that component from another, easily located component,
such as the nose, is determined in the following manner. An inner volume in which the engine
is expected is defined, and an outer volume surrounding the inner volume is also defined. The
LRF is directed at the inner and outer volumes to produce echoes from both volumes. A ratio
is taken of the number of echoes in the inner volumes to the number of echoes in both volumes.
If that echo exceeds a given threshold, the engine is determined to be present in the inner volume,
and the aircraft is considered to be identified. If the identification of the aircraft is still
ambiguous, another aircraft criterion, such as the tail, can be detected.
The aircraft criteria chosen for the second phase of the identification are physical
differences that can be detected by a laser range finder. An example of such a criterion is the
position, sideways and lengthwise, of an engine in relation to the aircraft nose. To consider an
aircraft identified, the echo pattern must not only reflect a fuselage of correct shape. It must also
reflect that there is an engine at a position, relative to the nose, where the expected aircraft does
have an engine. Other examples of criteria that can be used are the position of the main gear, the
position of the wings and the position of the tail.
The matching is preferably done only against the criteria specific for the expected aircraft
type. It would be very time consuming to match against the criteria of all other possible types.
Such matching would have to be against every type of aircraft that may land at a specific airport.
For each gate there is a defined a stopping position for each aircraft type that is planned
to dock at that gate. There might be a safety risk for any other type to approach the gate. The
stopping position is defined so that there is a sufficient safety margin between the gate and the
aircraft to avoid collision. The stopping position for each aircraft type is often defined as the
position of the nose gear when the door is in appropriate position in relation to the gate. There
is a database in the system where the distance from the nose to the nose gear is stored for each
aircraft type. The docking system guides the aircraft with respect to its nose position and stops
the aircraft with its nose in a position where the correct type will have its nose gear in the correct
stop position. If the wrong type is docked and if it has its wings or engines closer to the nose than
the correct type, there is a risk of collision with the gate.
During the aircraft criteria phase, all aircraft criteria specified for the expected aircraft
type can be checked. If an aircraft has a profile that can be used to discriminate it from any other
type, which is rarely the case, the profile will be the only aircraft criterion. Otherwise, another
criterion such as the position of the engine is checked, and if the identification is still ambiguous,
still another criterion such as the position of the tail is checked.
The LRF is directed to obtain echoes from the inner and outer volumes. If the ratio of
the number of echoes from within the inner volume to the number of echoes from within both
volumes is larger than a threshold value, the aircraft is identified as having an engine at the right
position, and that specific criterion is thus fulfilled. The ratio of the echo numbers is. however,
just an example of a test used to evaluate the presence of an engine at the right position or to
determine whether the echoes come from some other source, e.g., a wing. In cases in which that
is the only criterion, the aircraft is considered to be identified. Otherwise, the other specified
criteria (e.g., the height of the nose of the aircraft or evaluation of another aircraft criterion) have
to be fulfilled.
If necessary, several characteristics, such as the tail, gears, etc., can be used to identify
one specific type. The inner and outer volumes are then defined for each geometrical
characteristic to be used for the identification. The exact extension of the volumes is dependent
on the specific aircraft type and so is the threshold value.
A further identification criterion is the; nose height. The nose height is measured so as
to allow the horizontal scan to be placed over the tip of the nose. The measured nose height is
also compared with the height of the expected aircraft. If the two differ by more than 0.5 m, the
aircraft is considered to be of wrong type, and the docking is stopped. The value 0.5 m is given
by the fact that the ground height often varies along the path of the aircraft which makes it
difficult to measure with higher accuracy.
The invention lends itself to the use of "smart" algorithms which minimize the demand
on the signal processing at the same time as they minimize the effect of adverse weather and bad
reflectivity of aircraft surface. The advantage is that low-cost microcomputers can be used,
and/or computer capacity is freed for other tasks, and that docking is possible under almost all
weather conditions.
One important algorithm in that respect is the algorithm for handling of the reference
profiles. The profile information is stored as a set of profiles. Each profile in the set reflects the
expected echo pattern for the aircraft at a certain distance from the system. The position of an
aircraft is calculated by calculating the distance between the achieved echo pattern with the
closest reference profile. The distance interval between the profiles in the set is chosen so short
that the latter calculation can be made using approximations and still maintain the necessary
accuracy. Instead of using scaling with a number of multiplications, which is a demanding
operation, simple addition and subtraction can be used.
Another important algorithm is the algorithm for determining an aircraft"s lateral
deviation from its appropriate path. That algorithm uses mainly additions and subtractions and
only very few multiplications and divisions. The calculation is based on areas between the
reference profile and the echo pattern. As those areas are not so much affected by position
variationsor absence of individual echoes the algorithm becomes very insensitive to disturbances
due to adverse weather.
The calibration procedure enables a calibration check against an object at the side of the
system. The advantage is that such a calibration check is possible also when no fixed object is
available in front of the system. In most cases, there are no objects in front of the system that can
be used. It is very important to make a calibration check regularly. Something might happen to
the system, e.g., such that the aiming direction of the system is changed. That can be due to an
optical or mechanical error inside the system or it can be due to a misalignment caused by an
external force such as from a passing truck. If that happens, the system may guide an aircraft to
a collision with objects at the side of its appropriate path.
Another useful aspect of the present invention is that it can easily be adapted to take into
account the yaw angle of the aircraft. The yaw angle is useful to know for two reasons. First,
knowledge of the yaw angle facilitates accurate docking of the aircraft. Second, once the yaw
angle is determined, the profile is rotated accordingly, for more accurate matching.
In the verification process it is determined whether certain geometric characteristics, such
as an engine, are present in a certain position, e.g., relative to the nose. If the aircraft is directed
at an angle towards the docking guidance system (DGS), which is often the case, that angle has
to be known, in order to know where to look for the characteristics. The procedure is as follows:
1. Convert the polar coordinates (angle, distance) of the echoes to Cartesian coordinates
(x,y)-
2. Calculate the yaw angle.
3. Rotate the echo profile to match the yaw angle calculated for the aircraft.
4. Determine the existence of the ID characteristics.
The yaw angle is typically calculated through a technique which involves finding
regression angles on both sides of the nose of the aircraft. More broadly, the geometry of the part
of the aircraft just behind the nose is used. Doing so was previously considered to be impossible.
Still another aspect of the invention concerns the center lines painted in the docking ara.
Curved docking center lines are painted as the correct path for the nose wheel to follow, which
is not the path for the nose. If a DGS does not directly measure the actual position of the nose
wheel, the yaw angle is needed to calculate it based on measured data, such as the position of the
nose. The position of the nose wheel in relation to the curved center line can then be calculated.
BRIEF DESCRIPTION OF THE DRAWINGS
The features and advantages of the invention will become apparent from the following
detailed description taken in connection with the accompanying drawings wherein:
Fig 1 is a view illustrating the system as in use at an airport;
Fig 2 is a diagrammatic view illustrating the general componentry of a
preferred system in accordance with the present invention;
Fig 3 is a top plan view illustrating the detection area in front of a
docking gate which is established for purposes of detection and
identification of approaching aircraft;
Figs 4A and 4B together show a flow chart illustrating the main routine and the
docking mode of the system;
Fig 5 is a flow chart illustrating the calibration mode of the system;
Fig 6 is a view illustrating the components of the calibration mode;
Fig 7 is a flow chart illustrating the capture mode of the system;
Fig 8 is a flow chart illustrating the tracking phase of the system;
Fig 9 a is flow chart illustrating the height measuring the phase of the
system;
Fig 10 is a flow chart illustrating the identification phase of the system.
Fig 11 is a flow chart illustrating the aircraft criterion phase of the system;
Fig 12 is a diagram showing inner and outer volumes around an aircraft engine
used in the aircraft criterion phase;
Fig 13 is a diagram showing the tolerance limits of the measured nose-to-
engine distance for accepting an aircraft into a gate;
Fig 14 is a diagram showing the dependence of the safety margin on the
nose-to-engine distance in a situation in which an aircraft of the
wrong type is docked at the gate
Fig. 15 is a flow chart showing the basic steps used in recognizing an
aircraft which is at a yaw angle to the gate;
Fig. 15A is a diagram showing the geometry of the yaw angle;
Fig. 16 is a diagram showing the geometry used in determining the
regression lines -which are used in calculating the yaw angle;
Fig. 17 is a flow chart showing the steps used in calculating the yaw
angle;
Fig. 18 is a diagram showing the geometry used in rotating an echo
profile;
Fig. 19 is a flow chart showing the steps used in rotating the echo profile;
Fig. 20 is a flow chart showing the steps used in calculating an offset of
a nose wheel of an aircraft from a center line;
Fig. 21 is a diagram showing the geometry of the position of the nose
wheel relative to that of the nose; and
Fig. 22 is a diagram showing the geometry of the position of the nose
wheel relative to the center line.
Table I is a preferred embodiment of a Horizontal Reference Profile Table
which is employed to establish the identity of an aircraft in the
systems of the present invention;
Table II is a preferred embodiment of a Comparison Table which is
employed in the systems of the present invention for purposes of
effectively and efficiently docking an aircraft.
DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENT
Reference is now made to Figures 1-22 and Tables I-II, in which like numerals designate
like elements throughout the several views. Throughout the following detailed description,
numbered stages depicted in the illustrated flow diagrams are generally indicated by element
number in parenthesis following such references.
Referring to Fig. 1, the docking guidance systems of the present invention generally
designated 10 in the drawings provide for the computerized location of an object, verification of
the identity of the object and tracking of the object, the object preferably being an aircraft 12.
In operation, once the control tower 14 lands an aircraft 12, it informs the system that a plan is
approaching gate 16 and the type of aircraft (i.e., 747, L-1011, etc.) expected. The system 10
then scans the area in front of the gate 16 until it locates an object that it identifies as an airplane
12. The system 10 then compares the measured profile of the aircraft 12 with a reference profile
for the expected type of aircraft and evaluates other geometric criteria characteristic of the
expected aircraft type. If the located aircraft does not match the expected profile and the other
criteria, the system informs or signals the tower 14 and shuts down.
If the object is the expected aircraft 12, the system 10 tracks it into the gate 16 by .
displaying in real time to the pilot the distance remaining to the proper stopping point 29 and the
lateral position 31 of the plane 12. The lateral position 31 of the plane 12 is provided on a
display 18 allowing the pilot to correct the position of the plane to approach the gate 16 from the
correct angle. Once the airplane 12 is at its stopping point 53, that fact is shown on the display
18 and the pilot stops the plane.
Referring to Fig. 2, the system 10 includes a Laser Range Finder (LRF) 20, two mirrors
21, 22, a display unit 18, two step motors 24, 25, and a microprocessor 26. Suitable LRF
products for use herein are sold by Laser Atlanta Corporation and are capable of emitting laser
pulses and receiving the reflections of those pulses reflected off of distant objects and computing
the distance to those objects.
The system 10 is arranged such that there is a connection 28 between the serial port of
the LRF 20 and the microprocessor 26. Through that connection, the LRF 20 sends measurement
data approximately every 1/400th of a second to the microprocessor 26. The hardware
components generally designated 23 of the system 20 are controlled by the programmed
microprocessor 26. In addition, the microprocessor 26 feeds data to the display 18. As the
interface to the pilot, the display unit 18 is placed above the gate 16 to show the pilot how far the
plane is from its stopping point 29. the type of aircraft 30 the system believes is approaching and
the lateral location of the plane 31. Using thait display, the pilot can adjust the approach of the
plane 12 to the gate 16 to ensure the plane is on the correct angle to reach the gate. If the display
18 shows the wrong aircraft type 30, the pilot can abort the approach before any damage is done.
That double check ensures the safety of the passengers, plane and airport facilities because if the
system tries to dock a larger 747 at a gate where a 737 is expected, it likely will cause extensive
damage.
In addition to the display 18, the microprocessor 26 processes the data from LRF 20 and
controls the direction of the laser 20 through its connection 32 to the step motors 24, 25. The
step motors 24,25 are connected to the mirrors 21,22 and move them in response to instructions
from the microprocessor 26. Thus, by controlli ng the step motors 24,25, the microprocessor 26
can change the angle of the mirrors 21, 22 and aim the laser pulses from the LRF 20.
The mirrors 21, 22 aim the laser by reflecting the laser pulses outward over the tarmac
of the airport. In the preferred embodiment, the LRF 20 does not move. The scanning by the
laser is done with mirrors. One mirror 22 controls the horizontal angle of the laser while the
other mirror 21 controls the vertical angle. By activating the step motors 24, 25, the
microprocessor 26 controls the angle of the mirrors and thus the direction of the laser pulse.
The system 10 controls the horizon! al mirror 22 to achieve a continuous horizontal
scanning within a ±10 degree angle in approximately 0.1 degree angular steps which are
equivalent to 16 microsteps per step with the Escap EDM-453 step motor. One angular step is
taken for each reply from the reading unit, i.e., approximately every 2.5 ms. The vertical mirror
21 can be controlled to achieve a vertical scan between +20 and -30 degrees in approximately
0.1 degree angular steps with one step every 2.5 ms. The vertical mirror is used to scan vertically
when the nose height is being determined and when the aircraft 12 is being identified. During
the tracking mode, the vertical mirror 21 is continuously adjusted to keep the horizontal scan
tracking the nose tip of the aircraft 12.
Referring to Fig. 3, the system 10 divides the field in front of it by distance into three
parts. The farthest section, from about 50 meters out, is the capture zone 50. In that zone 50, the
system 10 detects the aircraft"s nose and makes a rough estimate of lateral and longitudinal
position of the aircraft 12. Inside the capture zone 50 is the identification area 51. In that area,
the system 10 checks the profile of the aircraft 12 against a stored profile 51. In that area, the
system 10 checks the profile of the aircraft 12 in that region, related to a predetermined line, on
the display 18. Finally, nearest to the LRF 20 is the display or tracking area 52. In the display
area 52, the system 10 displays the lateral and longitudinal position of the aircraft 12 relative to
the correct stopping position with its highest degree of accuracy. At the end of the display area
52 is the stopping point 53. At the stopping point 53, the aircraft will be in the correct position
at the gate 16.
In addition to the hardware and software, the system 10 maintains a database containing
reference profiles for any type of aircraft it might encounter. Within that database, the system
stores the profile for each aircraft type as a horizontal and vertical profile reflecting the expected
echo pattern for that type of aircraft.
Referring to Table I, the system maintains the horizontal profile in the form of a Table
I whose rows 40 are indexed by angular step and whose columns 41 are indexed by distance from
the stopping position for that type of aircraft. In addition to the indexed rows, the table contains
a row 42 providing the vertical angle to the nose of the plane at each distance from the LRF a
row 44 providing the form factor, k, for the profile and a row 45 providing the number of profile
values for each profile distance. The body 43 of the Table I contains expected distances for that
type of aircraft at various scanning angles and distances from the stopping point 53.
Theoretically, the 50 angular steps and the 50 distances to the stopping point 53 would
require a Table I containing 50 x 50, or 2500, entries. However, Table I will actually contain far
fewer entries because the profile will not expect a return from all angles at all distances. It is
expected that a typical table will actually contain between 500 and 1000 values. Well known
programming techniques provide methods of maintaining a partially full table without using the
memory required by a full table.
In addition to the horizontal profile, the system 10 maintains a vertical profile of each
type of aircraft. That profile is stored in the same manner as the horizontal profile, except that
its rows are indexed by angular steps in the vertical direction and its column index contains fewer
distances from the stopping position than the horizontal profile. The vertical profile requires
fewer columns because it is used only for identifying the aircraft 12 and for determining its nose
height, which take place at a defined range of distances from the LRF 20 in the identification area
51. Consequently, the vertical profile stores only the expected echoes in that range without
wasting data storage space on unneeded values.
The system 10 uses the previously described hardware and database to locate, identify
and track aircraft using the following procedures:
Referring to Figs. 4A and 4B, the software running on the microprocessor performs a
main routine containing subroutines for the calibration mode 60, capture mode 62 and docking
mode 400. The microprocessor first performs the calibration mode 60, then the capture mode
62 and then the docking mode 400. Once the aircraft 12 is docked, the program finishes. Those
modes are described in greater detail as follows:
Calibration Mode
To ensure system accuracy, the microprocessor 26 is programmed to calibrate itself in
accordance with the procedure illustrated in Fig. 5 before capturing an aircraft 12 and at various
intervals during tracking. Calibrating the system 10 ensures that the relationship between the
step motors 24,25 and the aiming direction is known. The length measuring ability of the LRF
20 is also checked.
Referring to Fig. 6, for calibration, the system 10 uses a square plate 66 with a known
position. The plate 66 is mounted 6 meters from the LRF 20 and at the same height as the LRF
20.
To calibrate, the system sets (a,p) to (0,0), causing the laser to be directed straight
forward. The vertical mirror 22 is then tilted such that the laser beam is directed backwards to
a rear or extra mirror 68 which redirects the beam to the calibration plate 66. (100) The
microprocessor 26 then uses the step motors 24. 25, to move the mirrors 21,22 until it finds the
center of the calibration plate 66. Once it finds the center of the calibration plate 66, thg
microprocessor 26 stores the angles (acp, ßcp) at that point and compares them to stored expected
angles. (102) The system 10 also compares the reported distance to the plate 66 center with a
stored expected value. (102) If the reported values do not match the stored values, the
microprocessor 26 changes the calibration constants, which determine the expected values, until
they do. (104, 106) However, if any of those values deviate too much from the values stored at
installation, an alarm is given. (108)
Capture Mode
Initially, the airport tower 14 notifies the system 10 to expect an incoming airplane 12
and the type of airplane to expect. That signal puts the software into a capture mode 62 as
outlined in Fig. 7. In capture mode 62, the microprocessor 26 uses the step motors 24, 25 to
direct the laser to scan the capture zone 50 horizontally for the plane 12. That horizontal scan
is done at a vertical angle corresponding to the height of the nose of the expected type of aircraft
at the midpoint of the capture zone 50.
To determine the correct height to scan, the microprocessor 26 computes the vertical
angle for the laser pulse as:
ßf=arctan[(H-h)//f]
where H = the height of the LRF 20 above the ground, h = the nose height of the expected
aircraft, and /f = the distance from the LRF 20 to the middle of the capture zone 50. That
equation results in a vertical angle for the mirror 21 that will enable the search to be at the correct
height at the middle of the capture zone 50 for the expected airplane 12.
Alternatively, the system 10 can store in the database values for P/for different types of
aircraft at a certain distance. However, storing ßr limits the flexibility of the system 10 because
it can capture an aircraft 12 only a single distance from the LRF 20
In the capture zone 50 and using that vertical angle, the microprocessor 26 directs the
laser to scan horizontally in pulses approximately 0.1 degree apart. The microprocessor 26 scans
horizontally by varying a. the horizontal angle from a center line starting from the LRF 20,
between ±amax, a value defined at installation. Typically, amax is set to 50 which, using 0.1 degree
pulses, is equivalent to 5 degrees and results in a 10 degree scan.
The release of the laser pulses results in echoes or reflections from objects in the capture
zone 50. The detection device of the LRF 20 captures the reflected pulses, computes the distance
to the object from the time between pulse transmission and receipt of the echo, and sends the
calculated distance value for each echo to the microprocessor 26. The micro processor 26 stores,
in separate registers in a data storage device, the total number of echoes or hits in each 1 degree
sector of the capture zone 50. (70) Because the pulses are generated in 0.1 degree intervals, up
to ten echoes can occur in each sector. The microprocessor 26 stores those hits in variables
entitled sa where a varies from 1 to 10 to reflect each one degree slice of the ten degree capture
zone 50.
In addition to storing the number of hits per sector, the microprocessor 26 stores, again
in a data storage device, the distance from the LRF 20 to the object for each hit or echo. Storing
the distance to each reflection requires a storage medium large enough to store up to ten hits in
each 1 degree of the capture zone 50 or up to 100 possible values. Because, in many cases, most
of the entries will be empty, well known programming techniques an reduce those storage
requirements below having 100 registers always allocated for those values.
Once that data is available for a scan, the microprocessor 26 computes the total number
of echoes, ST in the scan by summing the sa"s. The microprocessor 26 then computes SA/S, the
largest sum of echoes in three adjacent sectors. (72) In other words, SY is the largest sum of (Sa-1,,
Sa, Sa+1)-
Once it computes SMand SI the microprocessor 26 determines whether the echoes are
from an incoming airplane 12. If SMis not greater than 24, no airplane 12 has been found and
the microprocessor 26 returns to the beginning of the capture mode 62. If the largest sum of
echoes, SM is greater than 24 (74), a "possible" airplane 12 has been located. If a "possible"
airplane 12 has been located, the microprocessor checks if SA/S T is greater than 0.5 (76), or the
three adjacent sectors with the largest sum contain at least half of all the echoes received during
the scan.
If SM /SM is greater than 0.5, the microprocessor 26 calculates the location of the center
of the echo. (78, 82) The angular location of the center of the echo is calculated as:
a, = av + (Sa+1 - Sa-1)/(Sa-2 + Sa + Sa+1.,)
where Sa is the Sa that gave SM and txv is the angular sector that corresponds to that Sa.
The longitudinal position of the center of the echo is calculated as


where the /avi are the measured values, or distances to the object, for the pulses that returned an
echo from the sector av and where n is the total number of measured values in that sector. (78,
82) Because the largest possible number of measured values is ten, n must be less than or equal
to ten.
However, if S M/ST close range. If the cause is an aircraft at close range, that aircraft is probably positioned fairly
close to the centerline so it is assumed that at should be zero instead of the above calculated
value and that /, should be the mean distance given by the three middle sectors. (80) If the
distance distribution is too large, the microprocessor 26 has not found an airplane 12 and it
returns to the beginning of the capture mode 62. (81).
After calculating the position of the aircraft 12, the system 10 switches to"docking mode
400.
Docking Mode
The docking mode 400 illustrated in Figs. 4A and 4B includes four phases, the tracking
phase 84, the height measuring phase 86, the profile recognition phase 404, and the aircraft
criteria phase 408. In the tracking phase 84, the system 10 monitors the position of the incoming
aircraft 12 and provides the pilot with information about axial location 31 and distance from the
stopping point 53 of the plane through the display 18. The system 10 begins tracking the aircraft
12 by scanning horizontally.
Referring to Fig. 8, during the first scan in tracing phase, the microprocessor 26 directs
the LRF 20 to send out laser pulses in single angular steps, a or, preferably, at 0.1 degree
intervals between (a, - ap - 10) and (a, + ap + 10), where a, is determined during the capture
mode 62 as the angular position of the echo center and ap is the largest angular position in the
current profile column that contains distance values.
After the first scan, a is stepped back and forth with one step per received LRF value
between (as - ap - 10) and (as + ap + 10), where as is the angular position of the azimuth
determined during the previous scan.
During the tracking phase 84, the vertical angle ß is set to the level required for the
identified craft 12 at its current distance from the LRF 20 which is obtained from the reference
profile Table I. The current profile column is the column representing a position less than but
closer to /,.
The microprocessor 26 uses the distance from the stopping point 53 to find the vertical
angle for the airplane"s current distance on the profile Table I. During the first scan, the
distance, lo calculated during the capture mode 62, determines the appropriate column of the
profile Table I and thus the angle to the aircraft 12. For each subsequent scan, the
microprocessor 26 uses the ß in the column of the profile Table I reflecting the present distance
from the stopping point 53. (112)
Using the data from the scans and the data on the horizontal profile Table I, the
microprocessor 26 creates a Comparison Table II. The Comparison Table II is a two dimensional
" table with the number of the pulse, or angular step number, as the index 91, i, to the rows. Using
that index, the following information, represented as columns of the table, can be accessed for
each row: lt 92, the measured distance to the obj ect on that angular step; /k 93, the measured value
compensated for the skew caused by the displacement (equal to I, minus the quality sm, the total
displacement during the last scan, minus the quality i times sp, the average displacement during
each step in the last scan, i.e., lr(sm-isp)); di 94, the distance between the generated profile and
the reference profile (equal to x,y, the profile value for the corresponding angle at the profile
distance j minus Iki); a, 95, the distance the nose of the aircraft and the measuring equipment
(equal to rj50, the reference profile value at zero degrees, minus d,); ac 96, the estimated nose
distance after each step (equal to a^, the nose distance at the end of the last scan, minus the
quantity i times sp); ad, the difference between the estimated and measured nose distance (equal
to the absolute value of ^ minus a,.); and Note 97 which indicates the echoes that are likely
caused by an aircraft.
During the first scan in the tracking phase 84. the system 10 uses the horizontal profile
column representing an aircraft position, j, less than but closest to the value of/,. For each new
scan, the profile column whose value is less; than but closest to (a - sm ) is chosen where am is
the last measured distance to the aircraft 12 and sm is the aircraft"s displacement during the last
scan. Additionally, the values of the profile are shifted sideways by as to compensate for the
lateral position of the aircraft. (112)
During each scan, the microprocessor 26 also generates a Distance Distribution Table
(DDT). That table contains the distribution of a si value as they appear in the Comparison Table
II. Thus, the DDT has an entry representing the number of occurrences of each value of a; in the
Comparison Table II in 1 meter increments between 10 to 100 meters.
After every scan, the system 10 uses the DDT to calculate the average distance a,,,, to the
correct stopping point 53. The microprocessor 26 scans the data in the DDT to find the two
adjacent entries in the DDT for which the sum of their values is the largest. The microprocessor
26 then flags the Note 97 column in the Comparison Table II for each row containing an entry
for aj corresponding to either of the two DDT rows having the largest sum. (114)
The system 10 then determines the lateral deviation of offset. (116) The microprocessor
26 first sets:
2d = a,max - amin
where amax and amin are the highest and lowest a values for a continuous flagged block
of di values in the Comparison Table II. Additionally, the microprocessor 26 calculates:
for the upper half of the flagged dj in the block and:
___Y2= Sd1
for the lower half of the block. Using Y, and Y2 "a" 116 is calculated as:
a - k x (Y, - Y, )/d2
where k is given in the reference profile. If "a" exceeds a given value, preferably set to one, it
is assumed that there is a lateral deviation approximately equal to "a". The /i- column of the
Comparison Table II is then shifted "a" steps and the Comparison Table II is recalculated. The
process continues until "a" is smaller than an empirically established value, preferably one. The
total shift, as, of the I, column is considered equal to the lateral deviation or offset. (116) If the
lateral offset is larger than a predetermined value, preferably set to one, the profile is adjusted
sideways before the next scan. (118, 120)
After the lateral offset is checked, the microprocessor 26 provides the total sideways
adjustment of the profile, which corresponds to the lateral position 31 of the aircraft 12, on the
display 18. (122)
The microprocessor 26 next calculates the distance to the nose of the aircraft, am
an, = S(flagged aj)/N
where N is the total number of flagged a;. From a,,,, the microprocessor 26 can calculate the
distance from the plane 12 to the stopping point 53 by subtracting the distance from the LRF 20
to the stopping point 53 from the distance of the nose of the aircraft. (124)
Once it calculates the distance to the stopping point 53, the microprocessor 26 calculates
the average displacement during the last scan, sm. The displacement during the last scan is
calculated as:
where am-1, and a,,, belong to the last two scans. For the first scan in tracking phase 84, Sm is set
to O.
The average displacement during each step is calculated as:
Sp.Sm/P
where P is the total number of steps for the last scan cycle.
The microprocessor 26 "will inform the pilot of the distance to the stopping position 53
by displaying it on the display unit 18, 29. By displaying the distance to the stopping position
29, 53 after each scan, the pilot receives constantly updated information in real time about how
far the plane 12 is from stopping.
If the aircraft 12 is in the display area 52, both the lateral 31 and the longitudinal position
29 are provided on the display 18. (126, 128) Once the microprocessor 26 displays the position
of the aircraft 12, the tracking phase ends.
Once it completes the tracking phase, the microprocessor 26 verifies that tracking has not
been lost by checking that the total number of rows flagged divided by the total number of
measured values, or echoes, in the last scan is greater than 0.5. (83) In other words, if more that
50% of the echoes do not correspond to the reference profile, tracking is lost. If tracking is lost
and the aircraft 12 is greater than 12 meters from the stopping point, the system 10 returns to the
capture mode 62. (85) If tracking is lost and the aircraft 12 is less than or equal to 12 meters from
the stopping point 53, the system 10 turns on the stop sign to inform the pilot that it has lost
tracking. (85, 87)
If tracking is not lost, the microprocessor 26 determines if the nose height has been
determined. (13) If the height has not yet been determined, the microprocessor 26 enters the
height measuring phase 86. If the height has already been determined, the microprocessor 26
checks to see if the profile has been determined (402).
In the height measuring phase, illustrated in Fig. 9, the microprocessor 26 determines the
nose height by directing the LRF 20 to scan vertically. The nose height is used by the system
to ensure that the horizontal scans are made across the tip of the nose.
To check the nose height, the microprocessor 26 sets ß to a predetermined value ßmax and
then steps it down in 0.1 degree intervals once per received/reflected pulse until it reaches ßmin,
another predetermined value. ßmax and ßmax are set during installation and typically are -20 and
30 degrees respectively. After ß reaches ßmin the microprocessor 26 directs the step motors, 24,
25 up until it reaches ßmax. That vertical scanning is done with a set to as, the azimuth position
of the previous scan.
Using the measured aircraft distance, the microprocessor 26 selects the column in the
vertical profile table closest to the measured distance. (140) Using the data from the scan and
the data on the vertical profile table, the microprocessor 26 creates a comparison table shown
herein as Table II. Table II is a two dimensional table with the number of the pulse, or angular
step number, as an index 91, i, to the rows. Using that index, the following information,
represented as columns of the table, can be accessed for each row: I, 92, the measured distance
• to the object on that angular step, lsi 93, the measured value compensated for the skew caused by
the displacement (equal to li; minus the quantity Sm, the total displacement during the last scan,
minus the quantity i times sp, the average displacement during each step in the last scan), d1 94,
the distance between the generated profile and the reference profile (equal to rii, the profile value
for the corresponding angle at the profile distance j, minus /ki), a; 95. the distance between the
nose of the aircraft and the measuring equipment equal to rj50, the reference profile value at zero
degrees, minus dj), ac 96, the estimated nose distance after each step (equal to am, the nose
distance at the end of the last scan, minusthe jquantityi,times Sp),ad the.difference: between the
estimated and measured nose distance (equal to the absolute value of a1, minus ae), and Note 97
which indicates echoes that are likely caused by an aircraft 12.
During each scan the microprocessor 26 also generates a Distance Distribution Table
(DDT). That table contains the distribution of ai; values as they appear in Table II. Thus, the
DDT has an entry representing the number of occurrences of each value of ai in Table II in 1
meter increments between 10 to 100 meters.
After every scan, the system 10 uses the DDT to calculate the average distance, am, to the
correct stopping point 53. The microprocessor 26 scans the data in the DDT to find the two
adjacent entries in the DDT for which the sum of their values is the largest. The microprocessor
26 then flags the Note 97 column in Table II for each row containing an entry for a,-
corresponding to either of the two DDT rows having the largest sum. (142)
Once it completes the calculation of the average distance to the correct stopping point 53,
the microprocessor 26 calculates the average displacement during the last scan, sm. The
displacement during the last scan is calculated as:
sm = am-1 - am
where am-1 and am belong to the last two scans. For the first scan in tracking phase 84, sm is set
to 0. The average displacement sp during each step is calculated as:
sp = sm/P
where P is the total number of steps for the last scan cycle.
Calculating the actual nose height is done by adding the nominal nose height,
predetermined height of the expected aircraft when empty, to the vertical or height deviation.
Consequently, to determine the nose height, the system 10 first determines the vertical or height
deviation (144) vertical deviation is calculated by setting;
2d = ßmax - ßmin
where ßmax and ßmin are the highest and lowest ß value for a continuous flagged block of di; values
in the Comparison Table II. Additionally, the microprocessor 26 calculates:
Y1 = 14
for the upper half of the flagged di in the block and;
for the lower half of the block. Using Y, and Y2 , "a" is calculated as
a = k x (Y, - Y,)/d2
where k is given in the reference profile. If "a" exceeds a given value, preferably one, it is
assumed that there is a vertical deviation approximately equal to "a". The 1; column is then
shifted "a" steps, the Comparison Table II is re-screened and "a" recalculated. That process
continues until "a" is smaller than the given value, preferably one. The total shift, ßss of the li
column is considered equal to the height deviation. (144) The ßi values in the vertical
Comparison Table II are then adjusted as ßj + ?ßj where the height deviation ?ßj is:
?ßj = ßsx(aniß + as)/(aj + as)
and where amß is the valid am value when ßs was calculated.
Once the height deviation is determined, the microprocessor 26 checks if it is bigger than
a predetermined value, preferably one. (146) If the deviation is larger than that value, the
microprocessor 26 adjusts the profile vertically corresponding to that offset. (148) The
microprocessor 26 stores the vertical adjustment as the deviation from the nominal nose height.
(150) The actual height of the aircraft is the nominal nose height plus the deviation.
If the nose height is determined, or once the height measuring phase 86 is run, the
microprocessor 26 enters the identification phase illustrated in Fig. 10. (133, 88) In the:
identification phase 88, the microprocessor 26 creates a Comparison Table II to reflect the results
of another vertical scan and the contents of the profile table. (152, 154). Another vertical scan
is performed in the identification phase 88 because the previous scan may have provided
sufficient data for height determination but not enough for identification. In fact, several scans
may need to be done before a positive identification can be made. After calculating the vertical
offset 156, checking that it is not too large (158) and adjusting the profile vertically
corresponding to the offset (160) until the offset drops below a given amount, preferably one, the
microprocessor 26 calculates the average distance between marked echoes and the profile and
the mean distance between the marked echoes and that average distance. (162)
The average distance dm between the measured and corrected profile and the deviation
T from that average distance are calculated after vertical and horizontal scans as follows:
dm = Idi/N
T-Sldi- dm|/N
If T is less than a given value, preferably 5, for both profiles, the aircraft 12 is judged to be of the
correct type provided that a sufficient number of echoes are received. (164) Whether a sufficient
number of echoes is received is based on;
N/size > 0.75
where N is the number of "accepted" echoes and "size" is the maximum number of value:,
possible. If the aircraft 12 is not of the correct type, the microprocessor turns on the stop sign
136 and suspends the docking mode 400.
If the profile is determined (402), or once the profile determination phase is run (404),
the microprocessor 26 determines whether the aircraft criterion is determined (406). If not, the
aircraft criterion phase 408, which is illustrated in figs.l 1 and 12, is run.
In order for the criterion to be fulfilled, echoes must be returned from the location where
there is an engine on the expected aircraft. As there is some measurement uncertainty, there
might be echoes that actually come from the engine but appear to come from outside the engine.
Therefore, there must be defined a space Vi, called the inner volume or the active volume, around
the engine, such that echoes from within Vi are considered to come from the engine. Fig. 12
shows a sample Vi around an engine 13 of an airplane 12.
An engine is characterized in that for a horizontal scan there is a reflecting surface
surrounded by free space. In order to be able to discriminate between an engine and, e.g., a wing,
there must be defined another space Vo around the engine where there must be no or very few
echoes. The space Vo is called the outer volume or the passive volume. Fig. 12 also shows a
sample Vo around Vi.
The engine is defined by its coordinates (dx, dy, dz) for the center of the engine front
relative to the nose and by its diameter D. Those parameters are stored in a database for all
aircraft types.
Vi and Vo are defined by the extension sideways (x-direction) and lengthwise (z-
direction) from that engine center. The vertical position of the engine is given as (nose height +
dy).
For an engine on the wing, Vi and Vo are defined, by the following ranges of coordinates:
Vi:
x-direction: ±(D/2 + 1 m)
z-direction: + 3 m, - 1 m
Vo:
x-direction: ±2 m from Vi
z-direction: ±1.5 m from Vi
For tail engines the definition is the same except for Vo in the x-direction, which is given
by + 2 m from Vi. Otherwise echoes from the fuselage could fall within Vo and the
criterion would not be fulfilled.
Finally, the criterion is
Vi/(Vi + Vo) > 0.7
The threshold value 0.7 in the criterion is determined empirically. So are the limits given above
for Vi and Vo. At the moment those values are chosen so that unnecessary ID failures are
avoided and they are different only dependent on if the engine is on the wing or on the tail. As
docking data is accumulated they will be adjusted, probably different for different aircraft types,
to achieve better and better discrimination.
The aircraft criteria phase 408 applies the above principles as shown in the flow chart of
Fig. 11. When the aircraft criteria phase starts, the LRF is directed toward the engine or other
selected aircraft criterion in step 1102. In step 1104, the number of echoes in Vi is found, and
in step 1106, the number of echoes in Vo is found. In step 1108, it is determined whether
Vi/(Vi+Vo) exceeds the threshold value. If so, the aircraft criterion is indicated as met (OK) in
step 1110. Otherwise, the aircraft criterion is indicated as unmet (not OK) in step 1112.
If the aircraft criterion has been determined (406), or once the aircraft criterion phase is
complete (408), the microprocessor 26 determines whether the aircraft 12 has been identified.
(410). If the aircraft 12 has been identified, the microprocessor 26 checks whether the aircraft
12 has reached the stop position. (412). If the stop position is reached, the microprocessor 26
turns on the stop sign, whereupon the system 10 has completed the docking mode 400. (414)
If the aircraft 12 has not reached the stop position, the microprocessor 26 returns to the tracking
phase 84.
If the aircraft 12 is not identified, the microprocessor 26 checks whether the aircraft 12
is less than or equal to 12 meters from the stopping position 53. (416) If the aircraft 12 is not
more than 12 meters from the stopping position 53, the system 10 turns on the stop sign to
inform the pilot that the identification has failed. (418) After displaying the stop sign, the
system 10 shuts down.
If the aircraft 12 is more than 12 meters from the stopping point 53, the microprocessor
26 returns to the tracking phase 84.
In one possible implementation, the nominal distance (longitudinal and lateral) from the
nose to the engine is used as the aircraft criterion. In that implementation, docking is stopped
if the nose-to-engine distance, as measured in step 408, is more than two meters shorter than that
for the expected aircraft. If the difference is within two meters, it may still be possible to accept
an aircraft of the wrong type safely. In the latter case, if the safety margin between the engine
and a structure of the airport gate is three meters for the correct type of aircraft, the safety margin
for the other type of aircraft is still at least one meter. Tests have shown that the engine position
can be located to within about ± 1 meter and that the nose height can be determined to within ±
0.5 meter.
Fig. 13 shows the nominal nose to engine distance of an aircraft 12. The distance from
the aircraft"s nose to its engine 13 is of particular concern, since the engine 13 is in such a
position that misidentification can result in a collision between the engine 13 and a component
of the gate. Also shown are forward and backward tolerance limits for the position of the engine
13 that define the forward and backward extentS of Vi.
Fig. 34 shows an application of the identification procedure described above and in
particular shows what may happen if the system is set up for a selected aircraft 12A, but another
aircraft 12B attempts to dock at that gate. If a type of aircraft 12B different from the selected
aircraft 12A is accepted into the gate, the aircraft 12B will be stopped with the nose in the same
position in which the nose of the selected aircraft 12A would be stopped. As a result, the safety
margin, which is the distance from the engine to the closest component of the gate, such as the
bridge 15. is different between the aircraft 12A and 12B if the nose-to-engine distances of those
aircraft are different. As can be seen from Fig. 14, the safety margin for the aircraft 12B is equal
to the safety margin for the aircraft 12A minus the difference in nose-to-engine distances. If, for
example, the safety margin for the aircraft 12A is 3 m, and the nose-to-engine distance for the
aircraft 12B is 3.5m shorter than that for the aircraft 12A, the engine 13B of the aircraft 12B will
collide with the bridge 15. Therefore, if all aircraft types for which the nose-to-engine distance
is too small in comparison with that for the selected aircraft 12A are stopped, i.e., not accepted
into the gate, the safety margin can always be kept at an acceptable level.
A situation in which the aircraft is at an angle relative to the DGS 10 will now be
considered. As shown in Fig. 15A, a first aircraft 12D can be aligned correctly relative to the
DGS 10, whereas a second aircraft 12D can deviate from the correct alignment by a yaw angle
Y- A very high-level description of the technique used in such a situation is that the yaw angle
of the aircraft is determined, and the profile is rotated to match that yaw angle.
Fig. 15 shows a flow chart of the technique. In step 1502, the polar coordinates of the
echoes returned from the aircraft are converted to Cartesian coordinates. In step 1504, the yaw
angle is calculated. In step 1506, the echo profile is rotated. In step 1508, the ID characteristics
are detected in the manner already described.
Step 1502 is carried out in the following manner. The echo coordinates received from
the aircraft are converted from polar coordinates (afrf) to Cartesian coordinates (Xj,yj) with the
origin in the nose tip (anosernose) and with the y-axis along the line from the laser unit through the
nose tip as follows:
Xj - rj sinaj
yj = Tj COSa.J - rnose.
Step 1504 is carried out in a manner which will be explained with reference to Figs. 16
and 17. Fig. 16 is a diagram showing the geometry of the regression lines on either side of the
nose tip. Fig. 17 is a flow chart showing steps in the algorithm.
The algorithm is based on regression lines, calculated for echoes in a defined region
behind the nose tip. If there are a sufficient number of echoes on both sides of the nose, then the
yaw angle is calculated from the difference in angle between the regression lines. If only the
regression line for one side of the nose can be calculated, e.g. due to the yaw angle, then the yaw
angle is calculated from the difference in angle between that regression line and the
corresponding part of the reference profile.
In step 1702, the echo coordinates are converted to Cartesian coordinates (xfyj) in the
manner described above. In step 1704, the approximate coordinates of the nose tip are
calculated.
In step 1706, the echoes are screened in the following manner. Echoes not representative
for the general shape of the echo picture are removed before the angle of the echo picture is
calculated. The echo screening starts from the origin (the pointed out nose tip) and removes
both echoes if an echo at next higher angular step is at the same or shorter distance.
In step 1708, for each echo, the distance Rm tothe nose tip is calculated as
In step 1710, for each side of the nose tip, the echoes are selected for which R are larger
than Rmm, which is a constant (in the order of 1 - 2 m) defined specifically for each aircraft type.
In step 1712, the following mean values are calculated:
where n = the number of echoes > Rmin on a respective side, and the subscript right or left
identifies the respective side to which a particular quantity applies.
In step 1712, each regression line"s angle vreg to they axis is calculated as:
The subscript mean should be read as leftmean or rightmean in accordance with whether the
angle is calculated on the left or right side of the nose.
The. yaw angle y is calculated in the following manner. In step 1714, it is determined
whether the number n of echoes on both sides of the nose is greater than a predetermined value
TV, e.g., 5. If so, then in step 1718, y is calculated as
Y = (yreglefi+yreglefi)/2,
where yreglefi, and vregrighl are the angles calculated for the left and right sides of the nose using the
procedure of step 1712. On the other hand, if the n profile is used for the calculation. In step 1720, the side and segment of the profile are identified
which correspond to the side where n>N In step 1722, the angle yreglefi is calculated for that
segment using the procedure of step 1712. Then y is calculated in step 1718 as y = (yreglefi - vreg).
Once the yaw angle is calculated, then, in step 1506, the echo profile is rotated
accordingly. More specifically, the echo profile is converted from one Cartesian coordinate
system (x,y) to another (w,v) which has the same origin but is rotated by an angle equal to the yaw
angle y, as shown in Fig. 18. The rotation of the echo profile will now be described with
reference to Figs. 18 and 19.
In step 1902, the approximate coordinates of the nose tip are calculated. In step 1904,
the echo coordinates are converted from polar to Cartesian coordinates (xp yi) with the nose tip
as the origin of the coordinate system. The technique for doing so has been described above.
In step 1906, the echo coordinates are converted from the (x,y) coordinate system to the (w,v)
coordinate system, as shown in Fig. 18, through the following formulae:
ui, = xi, cos y + yt sin y;
vi, — -Xj sin y +yi cos y.
The echo coordinates as thus rotated are used to identify the aircraft in the manner
described above.
It will now be described how to set parameters defining center lines (CL"s), curved as
well as straight, with reference to Figs. 20-22. One docking system can handle several center .
lines with the technique to be described.
The CL is specified as a piecewise linear curve, where a,/ are coordinates (a — sideways,
/-lengthways) for the breaking points and are used as the defining parameters. The number of
coordinates used is chosen with respect to required positioning accuracy. A straight CL is thus
defined by the coordinates of two points (e.g. at the clip distance and at a stop position). The
number of coordinates required for a curved CL depends on its radius.
The microprocessor 26 is used in the CL setting mode of step 2002. in which the CI."s
arc mapped in the microprocessor. A CL to be defined is selected from a menu. One or more
calibration poles with known height and a top which is easily recognised in the calibration
picture arc placed on different positions on that CL. For each pole, the height of the pole is typed
in, and the top of the pole as appearing in the calibration picture is clicked. The a and /
coordinates for the pole are automatically entered in the table for that CL. The procedure is
repeated for each pole. The coordinates for the various poles arc ordered in the table by their /
values. The number of poles needed depends on the type of CL, with a straight CL needing only
two and a curved CL needing more.
The calculation of the offset of the nose from the nose wheel will now be discussed. The
CL is normally given as the ideal nosc-wheel track, but the guidance given to the aircraft is
normally based on the nose position. That means, in case of a curved CL, that either the CL
coordinates must be converted to nose-coordinates, or the nose position must be converted to
nose-wheel position. The latter is chosen, which means that the yaw angle (v,,u) of the aircraft is
determined in step 2004 in the manner described above.
The nosc-wheel position (awlv) is calculated in step 2006 as follows:
a- • a. + /„„ x sin vnu AL ? Lnw x cos O (in rad.)
l - 1n + lnw~ x cos vNN
where
an /m: measured position of the nose;
lnw: nosc-wheel distance; and
vrvi: estimated yaw angle of aircraft.
The offset of the no.se wheel from the CL is calculated in step 2008 as follows:
Offset = a, - a,r +¦ (/„ - /,)(aal,a)""(/../ - /.)
where
a, /, is the CL-coordinatc pair with /.-value just below /,,; and
a.,.,, I,., is the ("I.-coordinate pair with /,-value just above /„.
The calculations of step 2006 will now be explained with reference to Fig. 21, in which:
/„„: nosc-wheel distance
v: estimated yaw angle of aircraft
x: estimated sideways position of nose-wheel
a,.w an, + x/(/n +lnw /nw x cos v) (in rad.)
lwn /n + /nw x cos v
x — lmm x sin v
The calculations of step 2008 will now be explained with reference to Fig. 22, in which
Xo/y0 represents the estimated position of the nose wheel and xj y, represents the breaking points
in the picccwisc-lincar model of a curved CL. The "real" offset from the CL is the distance
measured in a rip.ht angle to the CL. An approximation of that distance is the distance measured
in a right angle to the laser beam from the docking system. That distance corresponds to the
value (xm - x0) in Fig. 22. As the absolute value of the offset not is important, that approximation
is used. From Fig. 22, it follows that
Offset - (xm - x0) = x1, - x0 + (y0 - y,)(x,,, -x,)/(y,., - y,).
While a preferred embodiment of the present invention has been set forth in detail above,
those skilled in the art will readily appreciate that other embodiments can be realized within the
scope of the invention. For example, while the aircraft criterion phase 408 is disclosed as using
the ratio Vi/(Vi +-Vo). the difference Vi-Vo could be used instead. Also, the specific numerical
ranges disclosed above should be considered to be illustrative rather than limiting. Those skilled
in the art will be able to derive other numerical ranges as needed to adapt the invention to other
models ofaircraft or to the .specific needs of various airports. Furthermore, while regression lines
are a useful technique for determining the yaw angle, any other technique can be used.
Therefore, the present invention should be construed as limited only by the appended claims.
WE CLAIM :
1. A system for determining whether a detected object is a known object, the
known object having a known profile and also having a known feature at a known
location, the system comprising :
projecting means for projecting light pulses onto the detected object;
collecting means for collecting light pulses reflected off the detected object
and for detecting a shape of the detected object in accordance with the light
pulses ;
comparing means for comparing the detected shape with a profile
corresponding to the known shape and for determining whether the detected
shape corresponds to the known shape ; and
identifying means for identifying whether the detected object is the known
object by determining whether the detected object has the known feature at the
known location.
2. The system as claimed in claim 1, wherein :
for the known object, an inner volume is defined so as to contain the
known feature, and an outer volume is defined so as not to contain the known
feature ;
the identifying means determines whether the detected object has the
known feature in the known location in accordance with a number of light pulses
reflected from within the inner volume and a number of light pulses reflected
from within the outer volume.
3. The system as claimed in claim 2, wherein the outer volume is defined to
surround the inner volume.
4. The system as claimed in claim 2, wherein the identifying means
determines whether the detected object has the known feature in the known
location in accordance with whether
Vi/(Vi + Vo) > T,
where :
Vi = the number of light pulses reflected from the inner volume ;
Vo = the number of light pulses reflected from the outer volume ; and
T = a predetermined threshold value.
5. The system as claimed in claim 4, wherein T = 0.7.
6. The system as claimed in claim 2, wherein the identifying means controls
the projecting means to project light pulses into the inner volume and the outer
volume.
7. The system as claimed in claim 1, wherein :
the known object comprises a nose with a known nose height; and
the identifying means identifies whether the detected object is the known
object by detecting a nose height of the detected object and comparing the
detected nose height to the known nose height.
8. The system as claimed in claim 7, wherein the identifying means
compares the detected nose height to the known nose height by taking a
difference between the detected nose height and the known nose height.
9. The system as claimed in claim 8, wherein the identifying means
identifies the detected object as the known object only if the difference is less
than or equal to a threshold difference.
10. The system as claimed in claim 9, wherein the threshold difference is
0.5m.
11. The system as claimed in claim 1, wherein the comparing means
determines a yaw angle of the detected object.
12. The system as claimed in claim 11, wherein the comparing means rotates
"the profile corresponding to the known shape by an angle equal to the yaw
angle.
13. A method for determining whether a detected object is a known object, the
known object having a known profile and also having a known feature at a known
location, the method comprising :
(a) projecting light pulses onto the detected object;
(b) collecting light pulses reflected off the detected object and for
detecting a shape of the detected object in accordance with the light pulses ;
(c) comparing the detected shape with a profile corresponding to the
known shape and for determining whether the detected shape corresponds to
the known shape ; and
(d) identifying whether the detected object is the known object by
determining whether the detected object has the known feature at the known
location.
14. The method as claimed in claim 13, wherein :
for the known object, an inner volume is defined so as to contain the
known feature, and an outer volume is defined so as not to contain the known
feature ;
said step of identifying comprises determining whether the detected object
has the known feature in the known location in accordance with a number of light
pulses reflected from within the inner volume and a number of light pulses
reflected from within the outer volume.
15. The method as claimed in claim 14, wherein the outer volume is defined
to surround the inner volume.
16. The method as claimed in claim 14, wherein said step of identifying
comprises determining whether the detected object has the known feature in the
known location in accordance with whether
Vi/(Vi + Vo) > T,
where :
Vi = the number of light pulses reflected from the inner volume ;
Vo = the number of light pulses reflected from the outer volume ; and
T = a predetermined threshold value.
17. The method as claimed in claim 16, wherein T= 0.7.
18. The method as claimed in claim 14, wherein said step of identifying
comprises controlling said step of projecting to project light pulses into the inner
volume and the outer volume.
19. The method as claimed in claim 13, wherein :
the known object comprises a nose with a known nose height; and
said step of identifying comprises identifying whether the detected object
is the known object by detecting a nose height of the detected object and
comparing the detected nose height to the known nose height.
20. The method as claimed in claim 19, wherein said step of identifying
comprises comparing the detected nose height to the known nose height by
taking a difference between the detected nose height and the known nose
height.
21. The method as claimed in claim 20, wherein said step of identifying
identifies the detected object as the known object only if the difference is less
than or equal to a threshold difference.
22. The method as claimed in claim 21, wherein the threshold difference is
0.5 m.
23. The method as claimed in claim 13, wherein said step of comparing
comprises determining a yaw angle of the detected object.
24. The method as claimed in claim 23, wherein said step of comparing
Comprises rotating the profile corresponding to the known shape by an angle
equal to the yaw angle.
25. A system for determining a yaw angle of a detected object, the system
comprising :
projecting means for projecting light pulses onto the detected object;
collecting means for collecting light pulses reflected off the detected object
and for detecting a shape of the detected object in accordance with the light
pulses ; and
angle determining means for determining the yaw angle from the shape
detected by the collecting means ;
wherein the detected object comprises a nose having a nose tip, and
wherein the angle determining means determines the yaw angle from a portion
of the shape which is adjacent to the nose tip.
26. The system as claimed in claim 25, wherein the nose has a left side and
a right side relative to the nose tip, and wherein the angle determining means
determines a regression line on at least one of the left side and the right side and
determines the yaw angle in accordance with the regression line.
27. A method of determining a yaw angle of a detected object, the method
comprising :
projecting light pulses onto the detected object;
collecting light pulses reflected off the detected object and detecting a
shape of the detected object in accordance with the light pulses ; and
determining the yaw angle from the shape detected by the collecting
means ; wherein the detected object comprises a nose having a nose tip, and
wherein said step of determining comprises determining the yaw angle from a
portion of the shape which is adjacent to the nose tip.
28. The method as claimed in claim 27, wherein the nose has a left side and a
right side relative to the nose tip, and wherein said step of determining comprises
determining a regression line on at least one of the left side and the right side
and determining the yaw angle in accordance with the regression line.
29. A system for determining whether a vehicle is following a center line, the
vehicle having a nose and a wheel, the system comprising :
a storage device for storing (i) coordinates representing a path of the
center line and (ii) a distance between the nose and the wheel ;
a detecting device for detecting (i) a position of the nose and (ii) a yaw
angle of the vehicle ; and
a calculating device for calculating (i) a position of the wheel, from the
position of the nose, the yaw angle detected by the detecting device and the
distance stored in the storage device, and (ii) an offset of the wheel from the
center line, from the coordinates stored in the storage device and the position of
the wheel.
30. A method of determining whether a vehicle is following a center line, the
vehicle having a nose and a wheel, the method comprising:
storing coordinates representing a path of the center line ;
storing a distance between the nose and the wheel ;
detecting a position of the nose ;
detecting a yaw angle of the vehicle ;
calculating a position of the wheel, from the position of the nose, the yaw
angle detected by the detecting device and the distance stored in the storage
device ; and
calculating an offset of the wheel from the center line, from the
coordinates stored in the storage device and the position of the wheel.
detecting a position of the nose ;
detecting a yaw angle of the vehicle ;
calculating a position of the wheel, from the position of the nose, the yaw
angle detected by the detecting device and the distance stored in the storage
device ; and
calculating an offset of the wheel from the center line, from the
coordinates stored in the storage device and the position of the wheel.
A laser range finder (LRF) (20) is used to identify an aircraft (12) approaching a gate. The LRF (20) is directed at
the aircraft (12), and from the echoes, a profile is derived and compared to known profiles. To distinguish among aircraft (12) with
similar profiles, the LRF is directed at a volume in which a feature such as an engine is expected and at another volume in which the
engine is not expected. The echoes from the two volumes are used to determine whether the engine is in its expected location. If so,
the aircraft (12) is identified as the correct type and is allowed to dock at the gate (16). Otherwise, the aircraft (12) is stopped. The
nose height can be used as yet another identifying criterion.

Documents:

IN-PCT-2002-544-KOL-CORRESPONDENCE.pdf

IN-PCT-2002-544-KOL-FORM 27.pdf

in-pct-2002-544-kol-granted-abstract.pdf

in-pct-2002-544-kol-granted-claims.pdf

in-pct-2002-544-kol-granted-correspondence.pdf

in-pct-2002-544-kol-granted-description (complete).pdf

in-pct-2002-544-kol-granted-drawings.pdf

in-pct-2002-544-kol-granted-examination report.pdf

in-pct-2002-544-kol-granted-form 1.pdf

in-pct-2002-544-kol-granted-form 18.pdf

in-pct-2002-544-kol-granted-form 2.pdf

in-pct-2002-544-kol-granted-form 3.pdf

in-pct-2002-544-kol-granted-form 5.pdf

in-pct-2002-544-kol-granted-gpa.pdf

in-pct-2002-544-kol-granted-letter patent.pdf

in-pct-2002-544-kol-granted-reply to examination report.pdf

in-pct-2002-544-kol-granted-specification.pdf

in-pct-2002-544-kol-granted-translated copy of priority document.pdf


Patent Number 213794
Indian Patent Application Number IN/PCT/2002/544/KOL
PG Journal Number 03/2008
Publication Date 18-Jan-2008
Grant Date 16-Jan-2008
Date of Filing 26-Apr-2002
Name of Patentee SAFEGATE INTERNATIONAL AB
Applicant Address 893 SOUTH MATLACK STREET, SUITE 220, WEST CHESTER, PA 19382
Inventors:
# Inventor's Name Inventor's Address
1 MILLGARD LARS BAGARVAGEN 3, S-831 52 OSTERSUND
PCT International Classification Number G 06 K 9/62
PCT International Application Number PCT/US00/29530
PCT International Filing date 2000-10-27
PCT Conventions:
# PCT Application Number Date of Convention Priority Country
1 09/429,609 1999-10-29 U.S.A.