DE212022000130U1 - Dynamic onboard system for automatically measuring the mass and balance, pitch, yaw, roll and equilibrium displacement of an aircraft on earth and in space - Google Patents

Abstract

Dynamic onboard method for automatically measuring the mass and balance, pitch angle, yaw, roll and balance displacement of an aircraft on the ground and in flight is the use of inertial navigation. The dynamic method consists of a body, accelerometers, inertial displacement sensors included in the composition, integrators and a calculator. The dynamic method is characterized by the fact that the calculations of the parameters mass and longitudinal centering as well as the pitch angle are carried out using differential measurements. O exist calculations before and after aircraft loading of the vertical displacements of the nose and tail sections of the fuselage. O there are calculations based on the signals from additionally installed accelerometers on the corresponding fuselage parts.

Description

The present invention relates to measurement and control systems and is intended to improve the flight safety and reliability of aircraft (LA), can be used for high-precision measurement of the spatial position, direction and movement of flight, as well as high-precision control of the trajectory of various dynamic objects.

The mass and balance of an aircraft are main indicators of controllability, stability, reliability and flight safety [1, 2, 3]. When loading and unloading the aircraft, moving cargo and passengers, dropping cargo, paratroopers, burning fuel, extending and retracting the landing gear, moving wing mechanizations and developing the aircraft, the position of the center of gravity is determined (CG) changed. Changes. During an aircraft's flight, the resulting external aerodynamic forces place stress on the aircraft and generate additional moments of force. To perform various maneuvers, it is necessary to fully or partially balance the external forces and moments acting on the aircraft relative to the alternating current of the aircraft. Therefore, it is necessary to know exactly the permissible range of mass and balance of the aircraft in order not to exceed it.

A report from the Netherlands Aerospace Laboratory from 1970 - 2005 shows that of the 82 accidents, 50 (61%) were civil aircraft accidents and 32 (39%) were cargo aircraft accidents due to exceeding the weight and center of gravity limits. In the reporting years 1970 - 2005, 93% of air traffic was passenger aircraft and 7% was cargo aircraft. Therefore, the risks associated with exceeding the weight and center of gravity limits for cargo aircraft exceed those for passenger aircraft by 8.5 times. In most cases, the causes of flight accidents are incorrect filling out of the balance sheet, incorrect load distribution in the front and rear luggage compartment, excessive take-off weight and load shifts. For the period 1998 - 2004, the document contains a large number of aviation safety reports from more than 40 operators, identifying 1200 aviation accidents (accidents) and analyzing some accidents (incidents) related to mass and balance [4].

The block diagrams of autopilots of modern aircraft do not provide for fault control that takes into account the influence of a dangerous deviation from the nominal value of the aircraft's balance [5, 6, 7, 8, 9].

A known method for determining the weight and position of the aircraft's center of gravity involves installing pressure sensors on the cylinder of each shock-absorbed landing gear, measuring the changing gas pressure in the cylinder cavity, and calculating the gas pressure based on averaging changes in gas pressure in the shock-absorbed cylinders of the Landing gear, forces acting on each landing gear, weight and position of the aircraft's center of gravity, measurement of gas pressure created when the aircraft rolls over the unevenness of the airfield, where conditions are created for multiple changes in gas pressure in the cylinders [10].

• der Einbau von Drucksensoren an den Federbeinen jedes Fahrgestells, was zusätzliche Befestigungspunkte der Sensoren mit sich bringt und zu einer verminderten Zuverlässigkeit der Federbeine des Fahrgestells führt;
• die Unfähigkeit, Roll-, Gier-, Nick- und Massenverschiebungen sowie den Schwerpunkt während des Fluges des Flugzeugs zu messen.

The disadvantages of the known method are:

• the installation of pressure sensors on the struts of each chassis, which entails additional mounting points of the sensors and leads to reduced reliability of the struts of the chassis;
• the inability to measure roll, yaw, pitch and mass shifts, as well as the center of gravity during aircraft flight.

There is a method for determining the mass of an aircraft and the location of its center of mass, implemented in a device in which the mass of the aircraft is converted into an electrical signal by summing at pressure sensors, for example piezoelectric sensors, the mass of the aircraft being determined from several sensors The position of the aircraft's center of mass is determined by distributing it to these sensors [11]. At the junction of the landing gear with the rest of the aircraft structure, pressure sensors are installed that record the pressure of its mass, which creates an electrical signal proportional to the pressure in the sensors.

The disadvantages of the known method also include the installation of additional attachment points for pressure sensors, which reduces the reliability of the shock-absorbed landing gear and makes it impossible to measure pitch, yaw, roll, mass displacement and centering in flight of the aircraft.

A known system for determining the position of the aircraft's center of gravity before takeoff includes a measuring weight platform (IP) installed in a square recess in the foundation on the taxiway (RD) of the runway (RWY). The IP is manufactured with fastening in the middle on hinged supports. The length of each arm of the IP is designed for simultaneous placement of the aircraft's front and main landing gear on it. A television broadcast camera and a camera are introduced into the system, the optical axes of which are aligned so that they are aligned on the MAR scale and intersect at a point above the MT rotation axis. A camera, the transmitting camera, is installed on the foundation of the platform to fix the MAR mark and mark the time of passage of the signals of the coordinates of the center of gravity. The output channel is connected to the radio line to convey information about the no-takeoff to the crew and ATC operators.

• Installation einer zusätzlichen Vertiefung des Fundaments auf dem Rollweg der Landebahn;
• bei der Messung der Koordinaten des Schwerpunkts des Flugzeugs wird zusätzliche Zeit benötigt, was zur Verzögerung der Flugleistung des Flugzeugs beiträgt,
• die Unmöglichkeit, Nick-, Gier-, Roll- und Massenverschiebungen sowie die Zentrierung im Flug des Flugzeugs zu messen.

The disadvantages of the known system are:

• Installation of an additional deepening of the foundation on the runway taxiway;
• when measuring the coordinates of the aircraft's center of gravity, additional time is required, which contributes to the delay in the flight performance of the aircraft,
• the inability to measure pitch, yaw, roll and mass displacements as well as centering in flight of the aircraft.

• - ein System bestehend aus: Kalibrierungsdatenspeicher, Speicher, einschließlich Reibungsdaten des Ausgangsteils der jeweiligen stoßdämpfenden Streben jedes Fahrwerks, gespeicherte Kalibrierungsdaten, einschließlich einer Vielzahl aufeinanderfolgender konstanter Auslassdrücke jedes Fahrwerks, die während des Be- und Entladens des Flugzeugs beobachtet werden, und einer Vielzahl von Auslassreibungsdaten, die dem jeweiligen Auslass zugeordnet sind.Dabei werden jeweilige Ausgangsdrücke und tatsächliche Belastung berechnet, die während des Kalibrierungsladens und -entladens auf die Beine wirkt.

An on-board system for determining the weight and balance of an aircraft [13] is known
has multiple landing gears, each of which includes a pneumatic shock absorption strut,

• - a system consisting of: calibration data memory, memory including friction data of the output part of the respective shock absorbing struts of each landing gear, stored calibration data including a plurality of successive constant outlet pressures of each landing gear observed during loading and unloading of the aircraft, and a plurality of outlet friction data , which are assigned to the respective outlet. This calculates the respective output pressures and actual load acting on the legs during calibration loading and unloading.

Pressure sensors are used to measure the pressure of the respective shock absorbers when the aircraft is subsequently loaded or unloaded; Load sensors are used to measure the vertical loads that act on the respective landing gear of the aircraft during subsequent loading or unloading.

Position sensors are also provided for measuring the position of the aircraft relative to the horizon during subsequent loading or unloading, as well as a computer for calculating the actual vertical load, which signals are received by each landing gear during subsequent loading or unloading.

The memory stores shock absorber outlet friction calibration data, shock absorber pressure read from the chassis load pressure sensors. A disadvantage of the known method is also the installation of additional attachment points for pressure sensors, which leads to a reduction in the reliability of the shock-absorbed chassis and the impossibility of measuring pitch, yaw, roll, mass displacement and alignment.

The closest technical approach to the claimed subject matter (method) is a method of using inertial navigation to determine the coordinates, roll angles, yaw, pitch and angular velocities of various aircraft [14].

The disadvantage of this method is the impossibility of measuring the mass and orientation as well as the displacement of the mass and orientation of the aircraft on the ground and in flight.

The system that technically comes closest to the object (system) in question is an inertial navigation system consisting of gyroscopes, accelerometers and correction blocks, which is intended to measure the coordinates, roll angles, pitch, yaw and angular velocities of the aircraft [14].

The disadvantage of this system is the impossibility of measuring the mass and balance of the aircraft as well as the displacement of the mass and the balance of the aircraft on the ground and in flight.

The aim of the invention is the autonomous measurement of the mass and orientation, displacement of the mass and orientation of the aircraft on the ground and in flight, as well as roll, yaw and pitch movements without the use of laser gyroscopes.

The technical result is achieved by installing additional accelerometers in the nose and tail at equidistant distances from the average of the center of gravity of the OO/aircraft along the Z axis on the wing tips, along the Y, X axes, parts that run along the corresponding axes are electrically connected to each other. As a result of loading or unloading, when the aircraft's center of gravity shifts, the nose and tail parts as well as the wing tips move unevenly relative to the average value of the OO/aircraft's center of gravity.

To measure the movement of the specified points of the aircraft, after opening the passenger doors, cargo compartments, lids of the filling tanks, additionally installed accelerometers are turned on to the perceived electrical signals of the position sensors (open and closed positions). The acceleration measured by these accelerometers, after double integration, determines the displacement of the corresponding points, after which the initial vertical positions of the nose and tail points of the fuselage (pitch channel) and the right and left parts of the wing tips (roll channel) are established. The vertical position signals are recorded in the memory of the calculator or computer of the corresponding channel. After closing the passenger doors, cargo compartments and tank cap, the additionally installed acceleration sensors are activated again by perceived electrical signals from the position sensors. In addition, after double integration of the accelerations obtained from additionally installed accelerometers, the final vertical positions of the nose and tail points of the fuselage (pitch channel) and the right and left parts of the wing tips (roll channel) are measured and established. These signals are recorded in the memory of the calculator or computer of the corresponding channel. The difference between the initial and final positions of the corresponding channels calculates the vertical displacements of the pitch channel and the roll channel. Thus, depending on the mass and centering of the aircraft, the vertical displacement of the tip and tail points of the fuselage, the right and left parts of the wing tips relative to the runway is determined. The computer determines the longitudinal centering and its displacement, the compensation arm and the actual pitch angle according to a specific algorithm that calculates the vertical displacements of the pitch channel. The computer calculates the vertical displacements of the roll channel using a certain algorithm and determines the transverse centering and its displacement, the actual roll angle. Before takeoff, the actual heading angle measured by other systems is entered into the yaw channel's computer memory as a reference signal. In the future, the current flight path of the aircraft can also be calculated in the field by measuring the yaw angle from the reference signal.

To achieve the goal, the corresponding accelerometers measure linear deviations relative to the center of gravity of the aircraft according to the differential scheme before and after loading by measuring the displacement of the extreme positions of the aircraft points on the ground and in flight. This helps determine relative to the X, Y and Z axes as well as roll angles, heading and pitch.

• • Trägheitsmessung der Roll-, Kurs- und Nickwinkel des Flugzeugs mit hoher Empfindlichkeit und Genauigkeit, jedoch ohne Lasergyroskop;
• • Erhöhung der Zuverlässigkeit und Zuverlässigkeit von Informationen;
• • Verbesserung der Flugsicherheit, die im Flugverkehr oberste Priorität hat;
• • Steigerung der Flugleistung.

The dynamic onboard method and system (device) for automatically measuring the mass and balance, pitch, yaw, roll and centering of an aircraft on the ground and in space offers a number of advantages:

• • Inertial measurement of aircraft roll, heading and pitch angles with high sensitivity and accuracy, but without a laser gyroscope;
• • Increase the reliability and reliability of information;
• • Improving aviation safety, which is a top priority in air travel;
• • Increase in flight performance.

The dynamic onboard method and system (device) for automatic measurement of mass and orientation, pitch, yaw, roll and offset of aircraft orientation on the ground and in space, which consists in the use of inertial navigation, are shown in the diagrams ( 1 - 6 ).

1 shows the arrangement of the longitudinal center of gravity using the example of the aircraft (AC) Boeing -747-8 [15] and the proposed arrangement of the accelerometers in it, which contributes to the implementation of an autonomous system (-board method for calculating the mass and balance in longitudinal direction direction, inclination, course and longitudinal displacement of aircraft in the middle on the ground and in flight).

2 shows the arrangement of the transverse center of gravity using the example of the Boeing -747-8 aircraft [15] and the proposed arrangement of accelerometers in it, which contributes to the implementation of an autonomous board (method for calculating the transverse mass and balance, roll and lateral displacement of the center of gravity the aircraft apparatus on the ground and in flight).

This method essentially places accelerometers at equal distances from the aircraft's average center of gravity along the Z axis at the tips of the side panels (wings) and along the Y, X axes front and back (nose and tail panels). From aircraft and sea vessels, it becomes possible to build an automatic system for measuring the deviation of the angular position of aircraft and sea vessels from a predetermined trajectory. Based on the proposed method, it is possible to introduce an additional autonomous automatic landing and mooring system on board.

wo
MAC (mittlere aerodynamische Sehne),
A - Abstand vom Koordinatenursprung zum Schwerpunkt (oder Masse) des Flugzeugs,
B - Abstand vom Koordinatenursprung zur MAR-Zehe,
C - Länge der mittleren aerodynamischen Sehne (MAC) des Flügels
sind.According to source [17], it is known that the center of gravity is defined as the ratio (A - B) of the distance from the center of gravity of the aircraft to the tip of the MAR (average aerodynamic chord) to the length. C of MAR in percent ( 3 ): $$\%MAC=\frac{A-\mathrm{<figure-callout\; id="100"\; label="b\; C"\; filenames="DE212022000130U1\_0047.png,DE212022000130U1\_0048.png"\; state="\{\{state\}\}">b</figure-callout>}}{C}100$$

where
MAC (mean aerodynamic chord),
A - distance from the origin of coordinates to the center of gravity (or mass) of the aircraft,
B - distance from the coordinate origin to the MAR toe,
C - Length of the wing's mean aerodynamic chord (MAC).
are.

4 shows a diagram of the position of the balance arm of the Boeing 747-8F aircraft when determining centering.

wobei BA (Balance Arm) der Abstand vom bedingten Koordinaten-ursprung (hinter der Nase des Flugzeugs in einem Abstand (1,78 m 70 Zoll)) zum Massenschwerpunkt ist. Der Abstand 1258 Zoll (31,95 m) ist der Abstand vom Koordinatenursprung bis zur Spitze des MAR, und der Abstand 327.8 Zoll (8,30 m) ist die Länge der mittleren aerodynamischen Flügelsehne. Der zulässige SAH-Bereich -%MAC =13-33% hängt von der Ausgleichsschulter BA ab (3). Aus Formel (1) und (2) ist es möglich, die Reichweite des Ausgleichsarms BA für ein bestimmtes Flugzeug zu bestimmen. $$A=BA=\frac{\%MAC\times C+B\times 100}{100}=\frac{\%MAC\times 327.8+1258\times 100}{100}$$

$$A=BA=1295÷1366.9\text{ oder }BA=32.893÷34.719\text{ M}$$

Using the example of the aircraft (aircraft) Boeing 747-8F ( 4 ) Centering (longitudinal centering) relative to the MAR is defined as [18]: $$\%MAC=\frac{bA-1258}{327.8}100$$

where BA (Balance Arm) is the distance from the conditional coordinate origin (behind the nose of the aircraft at a distance (1.78 m 70 inches)) to the center of mass. The distance 1258 inches (31.95 m) is the distance from the coordinate origin to the tip of the MAR, and the distance 327.8 inches (8.30 m) is the length of the center aerodynamic wing chord. The permissible SAH range -%MAC =13-33% depends on the compensation shoulder BA ( 3 ). From formula (1) and (2) it is possible to determine the range of the balancing arm BA for a specific aircraft. $$A=bA=\frac{\%MAC\times C+b\times 100}{100}=\frac{\%MAC\times 327.8+1258\times 100}{100}$$

$$A=bA=1295÷1366.9\text{or}bA=\mathrm{32,893}÷\mathrm{34,719}\text{M}$$

Taking a specific aircraft as an example, as can be seen from formulas (1) and (2), the variable parameter is the distance from the conditional origin of coordinates to the center of mass of the aircraft - BA, which depends on the loading or unloading of the aircraft. Formula (1) applies to many types of aircraft.

This method provides the necessary parameters for calculating the actual location of the aircraft's center of gravity and its displacement ( 1 ) determined. When the longitudinal center of gravity is shifted forward ΔL within the permissible range of the center of gravity (e.g. up to 11% MAR), the amount of vertical deflection of the bow A _V B _V decreases and the amount of vertical deflection of the stern increases $A_v^l{b}_v^l.$

If the center of gravity shifts to the stern part within the normal range (e.g. up to 33% MAR), the value of the vertical deflection of the bow increases. The value of the vertical deflection of the A _V D _V tail section drops to the value $A_v^l{D}_v^l.$

If additional accelerometers are installed on the longitudinal axis of the fuselage at equidistant distances from the average of the center of gravity of the OO / aircraft in the nose (at point A _V ) and in the tail, measurement of the vertical is possible.

Heckteil des Rumpfes.Based on the signals from these accelerometers, which integrate the accelerations twice, one obtains the vertical movements of nose A _V B _V and $A_v^l{b}_v^l$

Rear part of the fuselage.

In this case, it is possible to measure in real time the actual length of the balance arm, BA, which may also change during the flight (in fuel consumption, extension and retraction of the landing gear, wing and tail mechanization, movement of cargo, movement of passengers ). Based on this value, the actual orientation of the aircraft and its offset are determined.

(1).At a certain value of the pitch angle Δϑ, regardless of the shift in the center of gravity (front or rear center of gravity), the deviations of the values of the bow and stern change (according to the lever rule), but the sum remains constant: $A_v+b_v+A_v^l{b}_v^l=\text{const}$

( 1 ).

: $$\frac{A_v{B}_v}{A_{v}O}=\frac{A_v^l{B}_v^l}{A_v^{l}O}$$

sind.If according to 1 ) the bow and stern accelerometers installed at equal distances relative to the center of gravity, corresponding to the average value of the balancing shoulder, for example, they can be at a distance measuring the vertical displacement, according to formula $A=bA=L=A_{v}O=A_v^{l}O$

: $$\frac{A_v{b}_v}{A_{v}O}=\frac{A_v^l{b}_v^l}{A_v^{l}O}$$

are.

With a forward center shift due to a reduction in the length of the front arm and an increase in the length of the rear arm (lever rule) relative to the actual center of gravity of the aircraft: $$\frac{A_v{b}_v}{A_v^l{b}_v^l}<1$$

In the event of a shift in the aft center due to an increase in the forward part of the arm length and a decrease in the rear part of the arm length (lever rule) relative to the actual center of gravity of the aircraft: $$\frac{A_v{b}_v}{A_v^l{b}_v^l}>1$$

erhöht sich um ΔL, d. h. $L+\mathrm{\Delta }L=A_v^{l}O+\mathrm{\Delta }L=A_{v}O+\mathrm{\Delta }L=A_v^{l}C.$

When the center of gravity on the ΔL balance arm is shifted to the bow, it decreases by ΔL, ie A=BA=L-ΔL=A _v 0-ΔL=A _V C, and the length $A_v^{l}O$

increases by ΔL, ie $L+\mathrm{\Delta }L=A_v^{l}O+\mathrm{\Delta }L=A_{v}O+\mathrm{\Delta }L=A_v^{l}C.$

und die Länge $A_v^{l}O$

nimmt um ΔL, d. h. $L-\mathrm{\Delta }L=A_v^{l}O-\mathrm{\Delta }L=A_{v}O-\Delta L=A_v^l{C}^l$

ab.When the center of gravity is shifted to the rear area ΔL, the compensation arm increases by ΔL, ie $A=bA=L+\mathrm{\Delta }L=A_{v}O+\mathrm{\Delta }L=A_v{C}^l,$

and the length $A_v^{l}O$

decreases by ΔL, ie $L-\mathrm{\Delta }L=A_v^{l}O-\mathrm{\Delta }L=A_{v}O-\Delta L=A_v^l{C}^l$

away.

$$A_{v}O-\mathrm{\Delta }L=\frac{\left(A_{v}O+\mathrm{\Delta }L\right)A_v{B}_v}{A_v^l{B}_v^l}$$

oder $$L-\mathrm{\Delta }L=BA=\frac{\left(L+\mathrm{\Delta }L\right)A_v{B}_v}{A_v^l{B}_v^l}$$

In the case where the center of gravity is shifted to the bow, formula (4) can be written as: $$\frac{A_v{b}_v}{A_{v}O-\mathrm{\Delta }L}=\frac{A_v^l{b}_v^L}{A_{v}O+\mathrm{\Delta }L}$$

$$A_{v}O-\mathrm{\Delta }L=\frac{\left(A_{v}O+\mathrm{\Delta }L\right)A_v{b}_v}{A_v^l{b}_v^l}$$

or $$L-\mathrm{\Delta }L=bA=\frac{\left(L+\mathrm{\Delta }L\right)A_v{b}_v}{A_v^l{b}_v^l}$$

According to 1 The pitch angle can be determined: $$\vartheta =\text{arctg}\frac{L-\mathrm{\Delta }L}{A_v{b}_v}=\text{arctg}\frac{L+\mathrm{\Delta }L}{A_v^l{b}_v^l}$$

wo der einzige variable Parameter die Längszentrierungsverschiebung ΔL ist, von der der Ausgleichsarm, einschließlich der Längszentrierung des Flugzeugs, abhängt. Aus der Formel (7) wird der Längsversatz der Zentrierung ermittelt: $$\mathrm{\Delta }L=\frac{L\left(A_v^l{B}_v^l-A_v{B}_v\right)}{A_v^l{B}_v^l+A_v{B}_v}$$

Based on the proposed method, accelerometers measure the vertical movements of the nose A _V B _V and the tail parts of the fuselage $A_v^l{b}_v^l,$

where the only variable parameter is the longitudinal centering displacement ΔL, on which the balance arm, including the longitudinal centering of the aircraft, depends. The longitudinal offset of the centering is determined from formula (7): $$\mathrm{\Delta }L=\frac{L\left(A_v^l{b}_v^l-A_v{b}_v\right)}{A_v^l{b}_v^l+A_v{b}_v}$$

Through the proposed method, the location of the front or rear displacement of the longitudinal centering and its value are determined by the formulas (5), (6), (7), (9). From this it can be seen that these vertical movements of the fuselage nose and tail contribute to the dynamic determination of mass and centering using accelerometers installed equidistant (or proportionally) from the aircraft's center of gravity, located in the nose and tail of the aircraft ( 1 ).

(Bug- und Heckpunkte des Rumpfes), mit einer bekannten Rumpflänge von 76,25 m [18] ermittelt. Nehmen wir an, dass der Durchschnittswert des Schwerpunkts im Abstand BA =(32,893+34,719)/2=3 3 liegt (8m), gelten folgende Gleichungen: $$A_v{B}_v=A_v^l{B}_v^l=\frac{2\pi \times BA}{360°}\times \mathrm{\Delta }\vartheta$$

$$A_v{B}_v=A_v^l{B}_v^L=\frac{6.2832\times 33.8}{360°}\times 1°=0.59\text{M}$$

$$A_v{B}_v+A_v^l{B}_v^l=1.18\text{M}$$

Based on the proposed method, if the Boeing 747-8 F aircraft deviates from the current preset inclination by an angle, Δϑ = 1 ⁰ , the total vertical deviation $A_v{b}_v+A_v^l{b}_v^l$

(bow and stern points of the hull), with a known hull length of 76.25 m [18]. If we assume that the average value of the center of gravity is at the distance BA =(32.893+34.719)/2=3 3 (8m), the following equations apply: $$A_v{b}_v=A_v^l{b}_v^l=\frac{2\pi \times bA}{360°}\times \mathrm{\Delta }\vartheta$$

$$A_v{b}_v=A_v^l{b}_v^L=\frac{6.2832\times 33.8}{360°}\times 1°=0.59\text{M}$$

$$A_v{b}_v+A_v^l{b}_v^l=1.18\text{M}$$

Wenn der Abstand zwischen den Beschleunigungsmessern 28,647 m beträgt, dann ist $A_v{B}_v+A_v^l{B}_v^l=0.5\text{M}.$

If the accelerometers in the bow and stern areas are installed in a special way along the fuselage axis at equidistant distances equal to the distance from the conditional origin of coordinates to the average center of gravity, and the distance between the accelerometers is 57.295 m, then the total linear vertical deviation of the axis at the installation points of the accelerometer relative to the initial horizontal position when pitching with an angle Δϑ = 1 ^c equal $A_v{b}_v+A_v^l{b}_v^l=1\text{M}.$

If the distance between the accelerometers is 28.647 m, then is $A_v{b}_v+A_v^l{b}_v^l=0.5\text{M}.$

$$\frac{\mathrm{\Delta }{A}_v{B}_v}{A_v{B}_v}=\frac{{\vartheta }_g}{\mathrm{\Delta }\vartheta }$$

If accelerometers with a sensitivity threshold (lowest ability to measure acceleration) a _p = 10 ^-7 g (10 ^-6 м/ceк ² ) are used at an initial vertical speed of the aircraft of zero, then the sensitivity threshold for vertical movements is, ie the smallest Ability to measure vertical displacement in time t = 1 second, defined as $$\mathrm{\Delta }{A}_v{b}_v=\frac{{\alpha }_p{t}^2}{2}=\frac{{10}^{-6}\text{M/ce}{\mathrm{\kappa }}^2\times 1\text{ce}\mathrm{\kappa }}{2}=5\times {10}^{-7}\text{M}$$

$$\frac{\mathrm{\Delta }{A}_v{b}_v}{A_v{b}_v}=\frac{{\vartheta }_G}{\mathrm{\Delta }\vartheta }$$

Ausdruck (12), wird die Auflösung des Nickwinkelwerts, d. h. minimale vertikale Abweichung ϑ_g, entsprechend der Tonhöhenempfindlichkeitsschwelle bestimmt:

For values Δϑ = 1°, $\mathrm{\Delta }{A}_v{b}_v=A_v^l{b}_v^l=0.5\text{M},$

Expression (12), the resolution of the pitch angle value, that is, minimum vertical deviation ϑ _g , is determined according to the pitch sensitivity threshold:

If the theoretical trajectory of the straight line deviates from the flight altitude of the aircraft by an angle Δϑ = 1°, the length of which D = 10000км (if one point of the line is fixed), the displacement of the movable end point of this line from the initial position (i.e. the length of the large arc in height) with R = D = 10000км is defined by: $$\mathrm{\Delta }D=\frac{2\pi R}{360°}\times \mathrm{\Delta }\vartheta =\frac{6.2832\times 10000\text{KM}}{360°}\times 1°=\mathrm{174,533}\text{KM}$$

With a deviation by an angle ϑ _g = 0.000001°, the displacement of the end point of this line from the initial position along the vertical (the length of the small elevation arc) is defined as: $$\mathrm{\Delta }{d}_v=\frac{2\pi D}{360°}\times {\vartheta }_G=\frac{6.2832\times 10000\text{KM}}{360°}\times 0.000001°=0.1745\text{M}$$

By installing horizontal accelerometers in the front and aft parts of the fuselage at a distance of 28.647 m from the average center of gravity, the horizontal deflection of the fuselage along the course relative to the longitudinal line was measured at values of ΔΨ = 1°.

(15) definiert die Auflösung des Gierwinkelwerts, d. h. minimale horizontale Abweichung Ψ_g entsprechend der Gierempfindlichkeitsschwelle entlang des Kurses: $$\frac{\mathrm{\Delta }{A}_v{B}_v}{A_v{B}_v}=\frac{{\mathrm{\Psi }}_g}{\mathrm{\Delta \Psi }}$$

The formula $A_H{b}_H=A_H^l{b}_H^l=0.5\text{M}$

(15) defines the resolution of the yaw angle value, that is, the minimum horizontal deviation Ψ _g corresponding to the yaw sensitivity threshold along the course: $$\frac{\mathrm{\Delta }{A}_v{b}_v}{A_v{b}_v}=\frac{{\mathrm{\Psi }}_G}{\mathrm{\Delta \Psi }}$$

Based 1 The yaw angle can be determined on course: $$\mathrm{\Psi }=\text{arctg}\frac{L-\mathrm{\Delta }L}{A_H{b}_H}=\text{arctg}\frac{L+\mathrm{\Delta }L}{A_H^l{b}_H^l}$$

die Auflösung des Rollwinkelwertes, d. h. die minimale horizontale Auslenkung γ_g entsprechend der Rollempfindlichkeitsschwelle ist: $$\frac{\Delta A_V{B}_V}{A_V{B}_V}=\frac{Y_g}{\Delta \gamma }$$

After the accelerometers on the wing tips ( 2 ) were installed at a distance of 28.647 m from the average center of gravity to measure the lateral orientation and roll, with the values of Expression (17 Δγ = 1°) the $A_S{b}_S=A_S^l{b}_S^l=0.5\text{M}$

the resolution of the roll angle value, ie the minimum horizontal deflection γ _g corresponding to the roll sensitivity threshold, is: $$\frac{\Delta A_v{b}_v}{A_v{b}_v}=\frac{Y_G}{\Delta \gamma }$$

The transverse centering shift is determined analogously to (9) in the form: $$\Delta L_S=\frac{L_S\left(A_S^l{b}_S^{/}-A_S{b}_S\right)}{A_S^l{b}_S^{/}+A_S{b}_S}$$

According to 2 the roll angle can be determined: $$\mathrm{\gamma }=\text{arctg}\frac{L_S-\Delta L_S}{A_S{b}_S}=\text{arctg}\frac{L_S+\Delta L_S}{A_S^{/}{b}_S^{/}}$$

If the trajectory of the straight line of the aircraft's flight path deviates by an angle ΔΨ = 1 °, the length of which D = 10000км (if one point of the line is fixed) corresponds to the movement of the movable end point of this line from the initial position, that is, the length of a large Arc with radius R = D = 10000км is defined as $$\Delta D=\frac{2\pi R}{360°}\times \Delta \mathrm{\Psi =}\frac{6.2832\times 10000\text{KM}}{360°}\times 1°=\mathrm{174,533}\text{KM}$$

With a deviation by an angle Ψ _g = 0.000001°, the horizontal displacement of the end point of this line from the initial position (the length of the small arc) is defined as: $$\Delta d_H=\frac{2\pi D}{360°}\times {\mathrm{\Psi }}_G\mathrm{=}\frac{6.2832\times 10000\text{KM}}{360°}\times 0.000001°=0.1745\text{M}$$

Theoretical way to determine the flight path deviation of the Boeing 747-8 F aircraft

When flying at a distance of 10,000 km, the trajectory deviation can be 0.1745 m without using a laser gyroscope and alternative navigation data sources. Based on this method, it is possible to implement an additional high-precision automatic navigation, landing and mooring system.

Thus, the theoretical sensitivity of the proposed system is several orders of magnitude higher than that of satellite navigation and high-precision onboard laser gyroscopes used in modern aircraft and ships [14, 16].

In the initial starting position of the aircraft, as a reference signal, the measured flight heading is entered into the starting position with high accuracy through various existing methods, and then this information is stored in the computer memory. In addition, the proposed method can introduce a high-precision and high-sensitivity autonomous automatic navigation, landing or mooring system in the flight process relative to the reference heading signal.

The onboard dynamic system (device) for automatically measuring the mass and longitudinal orientation, pitch angle and longitudinal displacement of the aircraft on the ground and in flight, which consists in the application of the proposed method, is shown in the diagram ( 5 ).

• zwei Beschleunigungsmessern, deren Empfindlichkeitsachsen in der Nickebene ausgerichtet sind, die zur Messung der vertikalen Beschleunigung der Bug- und
• Heckpunkte des Rumpfes ausgelegt sind und an den Bug- und Heckpunkten auf derselben Längsachse des Rumpfes installiert sind,
• dem Rumpf,
• zwei Integratoren zur einfachen und doppelten Integration der vertikalen Beschleunigungen von Bug- und Heckpunkten,
• Rechner (Computereinheiten) des Verhältnisses der vertikalen Verschiebungen der Bug- und Heckpunkte,
• Nickwinkelrechner,
• Rechner für den Längsausrichtungsversatz,
• Ausgleichsarm-Rechner,
• Zentrierungsrechner relativ zur durchschnittlichen aerodynamischen Profiltiefe (MAC),
• vertikalem Flugbahncomputer.

The system consists of:

• two accelerometers, whose sensitivity axes are aligned in the pitch plane, which are used to measure the vertical acceleration of the bow and
• stern points of the hull are designed and installed at the bow and stern points on the same longitudinal axis of the hull,
• the hull,
• two integrators for single and double integration of the vertical accelerations of bow and stern points,
• Calculator (computer units) of the ratio of vertical displacements of the bow and stern points,
• pitch angle calculator,
• Longitudinal Alignment Offset Calculator,
• balance arm calculator,
• Centering calculator relative to average aerodynamic profile depth (MAC),
• vertical trajectory computer.

The vertical speed is determined with a simple integration of the vertical accelerations of the bow and stern points. Double integration is used to determine the vertical displacement of the corresponding points.

The onboard dynamic system (device) for automatic measurement of mass and longitudinal orientation, yaw angle and longitudinal offset of aircraft orientation on the ground and in flight, which consists in the application of the proposed method, is shown in the diagram ( 6 ).

• zwei Beschleunigungsmessern, deren Empfindlichkeitsachsen in der Gierebene ausgerichtet sind, die zur Messung der horizontalen Beschleunigung ausgelegt sind und senkrecht zur Längsachse an den Bug- und Heckpunkten des Rumpfes installiert sind,
• zwei Integratoren zur ein- und zweifachen Integration der horizontalen Beschleunigungen der Bug- und Heckpunkte,
• Rechner (Computereinheiten) des Verhältnisses der horizontalen Verschiebungen der Bug- und Heckpunkte,
• Gierwinkelrechner,
• Rechner für den Längsausrichtungsversatz,
• Ausgleichsarm-Rechner,
• Zentrierungsrechner relativ zur durchschnittlichen aerodynamischen Profiltiefe (MAC),
• Computer für horizontale Flugbahnen.

The system consists of:

• two accelerometers, whose sensitivity axes are aligned in the yaw plane, designed to measure horizontal acceleration and installed perpendicular to the longitudinal axis at the bow and stern points of the fuselage,
• two integrators for single and double integration of the horizontal accelerations of the bow and stern points,
• Calculator (computer units) of the ratio of horizontal displacements of the bow and stern points,
• yaw angle calculator,
• Longitudinal Alignment Offset Calculator,
• balance arm calculator,
• Centering calculator relative to average aerodynamic profile depth (MAC),
• Horizontal trajectory computer.

The horizontal speed is determined with a simple integration of the horizontal accelerations of the bow and stern points. The horizontal displacement of the corresponding points is determined using double integration.

The onboard dynamic system (device) for automatic measurement of mass and lateral orientation, roll angle and lateral displacement of aircraft orientation on the ground and in flight, which consists in the application of the proposed method, is shown in the diagram ( 7 ).

• zwei Beschleunigungsmessern, deren Empfindlichkeitsachsen in der Rollebene ausgerichtet sind, die dazu bestimmt sind, die vertikale Beschleunigung des rechten und linken Punktes der Flügelspitzen zu messen, und die an den entsprechenden Punkten des Flügels installiert sind,
• zwei Integratoren zur einfachen und doppelten Integration der Vertikalbeschleunigungen der rechten und linken Punkte der Flügelspitzen,
• Rechner (Rechnereinheiten) des Verhältnisses der vertikalen Verschiebungen der rechten und linken Punkte der Flügelspitzen,
• Bankwinkelrechner,
• vertikaler Flugbahncomputer.

The system consists of:

• two accelerometers, whose sensitivity axes are aligned in the roll plane, intended to measure the vertical acceleration of the right and left points of the wing tips and installed at the corresponding points of the wing,
• two integrators for single and double integration of the vertical accelerations of the right and left points of the wing tips,
• Calculator (calculator units) of the ratio of vertical displacements of the right and left points of the wing tips,
• bench angle calculator,
• vertical trajectory computer.

With a simple integration of the vertical accelerations of the right and left points of the wing tips, the vertical speed is determined. The vertical displacement of the corresponding wing points is determined with a double integration.

Similarly, the proposed method and system (device) can autonomously calculate the navigation and centering performance of underwater and surface ships and missiles with high accuracy.

SOURCES OF INFORMATION

1. 1. Principles of flying. YAAATPL. Handbook of theoretical knowledge. Oxford Aviation, Frankfurt, Germany, 2001. pp. 273-309 .
2. 2. David G Hull. Basics of aircraft flight mechanics. Springer Publishing. Berlin, Heidelberg, 2007. P. 298 .
3. 3. Weight and Balance Handbook. US Department of Transportation. Federal Aviation Administration. 2016. p.114 .
4. 4. Analysis of safety events related to aircraft weight and balance. 19th Annual European Aviation Safety Seminar (EASS), 12-14 March, Amsterdam, Netherlands, 2007, p. 23 . www.nlr.nl.
5. 5. Airbus A340-500/600. Technical Training Manual. Car flight. France, Airbus SAS, 2002, p. 117 .
6. 6. Boeing 747-8. Maintenance Training Manual. Autopilot flight director system. Boeing Company, 2016, p. 301 .
7. 7. Boeing 767. Aircraft Maintenance Manual. Car flight. The Boeing Company, USA, 2012. P. 701 .
8. 8th. Boeing 787. Laboratory notebook for training students. Autoflight and Thrust Management Systems: Avionics Book. Boeing Company, 2010, p. 235 .
9. 9. E-Jets (Embraer E-Jet 170/190). Maintenance Training Manual. Car flight. 2011.P.284 .
10. 10. Patent: RU 2319115 . IPC: G01G 19/07, C1. Mironov AD, Yunisov RR, Ivanov MT A method for determining the weight and location of the center of gravity of an aircraft. March 10, 2008, bull. No. 7.
11. 11. Patent: RU 2400405 . IPC: B64D 43/00. Kopylov GA, Kovalev VD A method for determining the mass of an aircraft, the location of its center of mass and a device for carrying it out. September 27, 2009, bull. No. 27.
12. 12. Patent: RU 2463567 . IPC: G01G 19/07, G01M 1/12. Berestov LM, Miroshnichenko L.Ya., Kalinin Yu.l., Makarova A.Yu., Teplova OV, Frolkina LV The system for determining the location of the center of gravity of the aircraft before takeoff. October 10, 2012, bull. No. 28.
13. 13. Patent: US 7,967,244 B2 . Onboard aircraft weight and balance system. Michael A Long _ Geoffrey E. Gouette. Jun 28, 2011.
14. 14. Laseref VI. Micro-inertial reference system. Honeywell Product Description, USA, Honeywell Inc., October 2016, p. 27 .
15. 15. Boeing 747-8. Maintenance Training Manual. Avionics (B2), EASA IR Part 66 (818) COMM2. Weight and balance system. 2016, pp. 382-433 .
16. 16. Barykin V.V., Kuleshov A.V., Podchezertsev V.P., Shcheglova N.N. Laser gyroscope. M.: MSTU named after N.E. Bauman, 2010, 23s.
17. 17. Mass and balance. YAAATPL. Handbook of theoretical knowledge. Germany, Frankfurt: Oxford Aviation, 2001, pp. 60-61 .
18. 18. Boeing 747-8F. Weight and balance. Control and charging manual. Boeing Company, 2016. P. 412

QUOTES INCLUDED IN THE DESCRIPTION

This list of documents listed by the applicant was generated automatically and is included solely for the better information of the reader. The list is not part of the German patent or utility model application. The DPMA assumes no liability for any errors or omissions.

Cited patent literature

• RU 2319115 [0074]
• RU 2400405 [0074]
• RU 2463567 [0074]
• US 7967244 B2 [0074]

Non-patent literature cited

• Principles of flying. YAAATPL. Handbook of theoretical knowledge. Oxford Aviation, Frankfurt, Germany, 2001. pp. 273-309 [0074]
• David G Hull. Basics of aircraft flight mechanics. Springer Publishing. Berlin, Heidelberg, 2007. P. 298 [0074]
• Weight and Balance Handbook. US Department of Transportation. Federal Aviation Administration. 2016. p.114 [0074]
• Analysis of safety events related to aircraft weight and balance. 19th Annual European Aviation Safety Seminar (EASS), 12-14 March, Amsterdam, Netherlands, 2007, p. 23 [0074]
• Airbus A340-500/600. Technical Training Manual. Car flight. France, Airbus SAS, 2002, p. 117 [0074]
• Boeing 747-8. Maintenance Training Manual. Autopilot flight director system. Boeing Company, 2016, p. 301 [0074]
• Boeing 767. Aircraft Maintenance Manual. Car flight. The Boeing Company, USA, 2012. P. 701 [0074]
• Boeing 787. Laboratory notebook for training students. Autoflight and Thrust Management Systems: Avionics Book. Boeing Company, 2010, p. 235 [0074]
• E-Jets (Embraer E-Jet 170/190). Maintenance Training Manual. Car flight. 2011.P.284 [0074]
• Laseref VI. Micro-inertial reference system. Honeywell Product Description, USA, Honeywell Inc., October 2016, p. 27 [0074]
• Boeing 747-8. Maintenance Training Manual. Avionics (B2), EASA IR Part 66 (818) COMM2. Weight and balance system. 2016, pp. 382-433 [0074]
• mass and balance. YAAATPL. Handbook of theoretical knowledge. Germany, Frankfurt: Oxford Aviation, 2001, pp. 60-61 [0074]
• Boeing 747-8F. Weight and balance. Control and charging manual. Boeing Company, 2016. p. 412 [0074]

Claims (6)

Dynamic onboard method for automatically measuring the mass and balance, pitch angle, yaw, roll and balance displacement of an aircraft on the ground and in flight is the use of inertial navigation. The dynamic method consists of a body, accelerometers, inertial displacement sensors included in the composition, integrators and a calculator. The dynamic method is characterized by the fact that the calculations of the parameters mass and longitudinal centering as well as the pitch angle are carried out using differential measurements. O exist calculations before and after aircraft loading of the vertical displacements of the nose and tail sections of the fuselage. O there are calculations based on the signals from additionally installed accelerometers on the corresponding fuselage parts. A dynamic airborne method for automatically measuring the mass and balance, pitch angle, yaw, roll and balance displacement of an aircraft on the ground and in flight is the use of inertial navigation. The dynamic method consists of a body, accelerometers, inertial displacement sensors included in the composition, integrators and a calculator. The dynamic method is characterized by the fact that the calculation of the parameters mass and longitudinal centering as well as the yaw angle along the course is carried out by differential measurements of the horizontal displacements of the nose and tail of the fuselage. O there are calculations based on the signals from additionally installed accelerometers on the corresponding fuselage parts. A dynamic airborne method for automatically measuring the mass and balance, pitch angle, yaw, roll and balance displacement of an aircraft on the ground and in flight is the use of inertial navigation. The dynamic method consists of a body, accelerometers, inertial displacement sensors included in the composition, integrators and a calculator. The dynamic method is characterized by the fact that the calculations of the parameters mass and transverse centering as well as the heeling angle are carried out using differential measurements. O there are calculations before and after loading the aircraft vertical movements of the tips of the right and left wing parts. O there are calculations based on the signals from additionally installed accelerometers on the corresponding parts of the wing. The dynamic on-board system (device) for automatically measuring the mass and balance, pitch angle, yaw, roll and balance displacement of an aircraft on the ground and in flight consists of an accelerometer, an integrator, a computer. The dynamic system is characterized by the fact that it is additionally equipped with vertical accelerometers for bow and stern, integrators for the vertical acceleration of bow and stern, calculators for the vertical displacement ratio of the bow and stern points, an inclination calculator, a calculator for the longitudinal centering offset, a Calculator for the balancing arm and a relative centering is equipped means aerodynamic chord calculator, a computer for vertical trajectories. The dynamic on-board system (device) for automatically measuring the mass and balance, pitch angle, yaw, roll and balance displacement of an aircraft on the ground and in flight consists of an accelerometer, an integrator, a computer. The dynamic system is different in the fact that it is additionally equipped with horizontal bow and stern accelerometers, integrators of horizontal bow and stern accelerations, calculators for the ratio of horizontal displacements of the bow and stern points, a yaw calculator along the course, a calculator for the longitudinal centering offset is equipped, a balancing arm calculator, a centering calculator relative to the average aerodynamic chord, a horizontal calculator for flight paths. The dynamic on-board system (device) for automatically measuring the mass and balance, pitch angle, yaw, roll and balance displacement of an aircraft on the ground and in flight consists of an accelerometer, an integrator, a computer. The dynamic system is characterized by the fact that the left and right parts of the wing tips are additionally equipped with vertical accelerometers, vertical acceleration integrators of the corresponding accelerometers, calculators for the ratio of vertical displacements of the left and right parts of the wing tips, etc. Roll calculator and a calculator for the transverse centering displacement.

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