IL134188A - Steering system and method - Google Patents

Description

iy& <))> ini nu>tn nmya Flying Object Steering System and Method Israel Aircraft Industries Ltd. 00II9I-IL FLYING OBJECT STEERING SYSTEM AND METHOD FIELD OF THE INVENTION The present invention relates to aerodynamic systems, in general and to systems and methods for steering an aircraft, in particular.

BACKGROUND OF THE INVENTION Systems for steering an aircraft are known in the art. In general, an aircraft, such as an airplane, a missile or a bomb, has a plurality of steering elements. Such elements are either dynamic fins or dynamic power nozzles. Conventional steering systems, which are designed for ballistic aircraft, employ a plurality of fins. Some of these systems utilize stationary fins, which are fixed to the body of the ballistic aircraft. The rest of these systems utilize adjustable fins, controlled in various ways.

US Patent No. 5,322,243 issued to Stoy, is directed to separately banking maneuvering, aerodynamic control surfaces, system and method. This reference discloses means for forming a forward section of the vehicle, which is free to rotate about the longitudinal axis independently of the roll attitude. Thus, the portion of the rotatable section supporting the pair of control fins is rotatable about the longitudinal axis of the flight vehicle. An internal control system inside the flight vehicle steers the vehicle by independently changing the attitude of the control surfaces about lateral axes to control the roll attitude of the section and to obtain a desired change in direction of flight. The system, according to this reference, includes actuators connected to the control surfaces for changing the relative attitude of the control surfaces by differential deflection, in which the surfaces move counter to each other, and by symmetrical deflection in which the surfaces move in the same direction, the attitudes of these surfaces being actuated independently so that differential deflection controls the rotation orientation of the section, and symmetrical deflection produces a pitching moment to the vehicle in the direction of the collective surface orientation. The flight vehicle has a fuselage or main body and conventional aerodynamic control surfaces located at the middle or rear of the aircraft, including a vertical stabilizer and a pair of wings, having movable control surfaces or ailerons.

US Patent No. 5,085,381 to Spiroff et al., is directed to a deployable aerodynamic aerosurface. This reference illustrates the aft section of a guided missile, which is provided with two fixed aerosurfaces and two pivotal aerosurfaces, also referred to as fins. The deployable fins are pivotally mounted about axes extending lengthwise of missile at opposite sides thereof between a folded or stowed position, and an unfolded or deployed extended flight position. Each fin comprises a root or base member and a control surface movable relative to the base member about an axis normal to the axis about which the base member pivots. The base member is suitably rigidly secured to a rotatable housing forming a part of the missile faring and nestled between faring fixed portions.

US patent no. 5,584,448 to Epstein et al., is directed to a flight control device. This reference describes a base, which is attached to a rocket body and is generally common to all the fins and the control devices associated therewith. The base includes means for supporting four fins and the mechanisms for actuating them. For each fin, there are provided two partitions. The four fins are folded before the missile is launched. They open wide, after launch and each of them can be adjusted to a specified position, so as to navigate the missile to its target.

US patent no. 4,667,899 to Wedertz, is directed to a double swing wing self-erecting missile wing structure. The missile air frame includes a nose portion, a tail portion, and a middle portion. The missile includes four self-erecting wing structures. The wing structures include attaching means for pivotally attaching the wing to the missile air frame in an erected position such as that illustrated. These attaching means include rotational components that enable rotation of the wing about its spanwise axis substantially and the simultaneous pivoting of the wing forward about an axis that is generally perpendicular to the spanwise axis, to a retracted position alongside the air frame. Attaching means includes means for enabling passing air to swing the wing to an erected position when it is released during flight. The rotational components serve this function, and erection is accomplished by simply releasing the wing from the retracted position.

US patent no. 5,326,049 to Rom et al., is directed to a device including a body having folded appendage to be deployed upon acceleration. The missile has four foldable wings. Each of the four foldable wings is mounted to the missile by a pivotal assembly. As soon as the missile leaves the canister and accelerates in the direction of its longitudinal axis, the acceleration of the missile produces an inertial force through the center of gravity CG opposite to the direction of acceleration of the missile. The reaction force produced at the center of gravity CG of the wing, multiplied by the distance "b" produces a moment pivoting the wing about pivot axis PP^. This pivoting of the wing about pivot axis PPi moves the wing center of gravity CG outwardly of the pivot axis PP2, so that the inertial force of the center of gravity CG of the wing now produces a moment tending to pivot the wing about pivot axis PP2. Thus, as soon as the missile leaves the canister, the wing tends to pivot about both pivot axes PPi and PP2.

US patent no. 4,964,593 to Kranz, is directed to a missile having rotor ring. The rotor ring has at least one adjustable fin actuated by the relative rotation between the rotor ring and the missile. At least one motor is arranged between the rotor ring and the missile, and a control device is provided which allows the motor to work as a generator and perform the fin adjustment.

Reference is now made to Figures 1A, 1 B and 1C. Figure 1A is an illustration of a conventional steering system, which is mounted on a missile, generally referenced 10. Figure 1 B is an illustration of the missile 10, of Figure 1A, without the steering system. Figure 1C is a view from the rear of the steering system of Figure 1A. With reference to Figure 1A, the steering mechanism, which is attached to missile 10 includes an avionics unit 12 and a steering system 24. The steering system 24 includes a plurality of fins, generally referenced 26A, 26B, 26C and 26D and is located near the rear end of the missile 10. With reference to Figure 1 B, the missile 10 has a special groove 16, which is located on the missile body 14, near the rear end 18 and is designed to receive a steering system such as system 24.

With reference to Figure 1C, steering system 24 includes two sections 24A and 24B, which are joined together around groove 16 (Figure 1 B). Section 24A includes fins 26A and 26B. Section 24B includes fins 26C and 26D. Steering system 24 is connected to the avionics unit 12 by means of communication conveyer 22. It is noted that this design implicates disadvantages in the assembly stage, in which the steering system is mounted on the missile. Such disadvantages include performing mechanical blind connections between sections 24A and 24B in groove 16, the fact that the steering system at the rear end of the missile is separated from the avionics at the front end of the missile.

SUMMARY OF THE PRESENT INVENTION It is an object of the present invention to provide a novel method for maneuvering a flying object, which overcomes the disadvantages of the prior art.

In accordance with the present invention, there is thus provided an apparatus for controlling a flight path of a flying object where the object includes at least one static fin. The apparatus includes a housing and at least two dynamic fins, rotatably connected to the housing. The static fin and the dynamic fins are substantially located on the same plane, which is normal to a longitudinal axis of the flying object. The flying object can be a self-propelled object or an unpropelled object.

The housing can be fastened to the object, within a predetermined groove on the object, where it completes an aerodynamic overall shape in combination with the flying object. The housing can contain at least one avionics unit therein, for actuating the dynamic fins.

In accordance with another preferred embodiment of the invention, there is thus provided a method for controlling a flight path of a flying object, which includes at least two dynamic fins and at least one static fin. The method includes the steps of: activating at least two of the dynamic fins such that the object substantially rolls about a longitudinal axis of the object, and activating at least two of the dynamic fins, thereby directing the object in a predetermined direction.

In accordance with a further preferred embodiment of the invention, there is provided a method which includes the steps of: detecting the orientation of the object in the flight path, determining a change in the orientation, and rotating the dynamic fins according to the change. This method can further include the step of determining a desired orientation. The method can also include the step of determining a required orientation change, according to the desired orientation and the orientation. Accordingly, the required orientation change includes a pitch change, a yaw change and a combination thereof.

The method can further include the step of determining a summation vector, which is the vectorial sum of a pitch axis and a yaw axis, determined according to the required orientation change. In addition, the method can also include the step of determining a direction and value of roll, according to the summation vector, an angle in direction of the summation vector, and the like.

Finally, the method can include the steps of rolling and directing the object according to the direction of roll, the value of roll, and the angle, by rotating the dynamic fins.

In accordance with another preferred embodiment of the present invention, there is thus provided a method for increasing the range of a flying object in at least one stage. The method includes the steps of detecting a turning point of the object, and directing the object toward a maximal point.

BRIEF DESCRIPTION OF THE DRAWINGS The present invention will be understood and appreciated more fully from the following detailed description taken in conjunction with the drawings in which: Figure 1A is an illustration of a conventional steering system, which is mounted on a missile, which is known in the art, as described herein above; Figure 1 B is an illustration of the missile of Figure 1A, without the steering system, as described herein above; Figure 1 C is a view from the rear of the steering system of Figure 1 A, as described herein above; Figure 2A is a schematic illustration of a ballistic aircraft, constructed and operative in accordance with a preferred embodiment of the present invention; Figure 2B is a schematic illustration of a steering system, for mounting on the ballistic aircraft of Figure 2A, constructed and operative in accordance with a preferred embodiment of the present invention; Figure 3 is an illustration in detail of the steering system of Figure 2B; Figure 4A is an illustration of the steering system of Figure 2B, at one mode of steering, operative in accordance with another preferred embodiment of the present invention; Figure 4B is an illustration of the steering system of Figure 2B, at another mode of steering, operative in accordance with another preferred embodiment of the present invention; Figure 4C is a schematic view from the rear of the steering system of Figure 2B, with the major steering parameters denoted thereon; Figure 5 is a schematic illustration of a method for operating the steering system, according to a further preferred embodiment of the present invention; Figure 6 is a schematic illustration of the navigation avionics unit, of the steering system of Figure 2B, constructed and operative in accordance with another preferred embodiment of the present invention; Figure 7A is an illustration in perspective of a steering system, constructed and operative in accordance with another preferred embodiment of the present invention; Figure 7B is a view from the rear of the system of Figure 7A; Figure 8 is a schematic illustration of a steering system, constructed and operative in accordance with a further preferred embodiment of the present invention; Figure 9, is an illustration of side view of a flight path of a missile, according to a preferred embodiment of the present invention; and Figure 10, is an illustration of side view of a flight path of a missile, according to another preferred embodiment of the present invention.

DETAILED DESCRIPTION OF PREFERRED EMBODIMENTS The present invention overcomes the disadvantages of the prior art by providing a novel steering system, having a plurality of fins, where some of the fins are dynamic and others are stationary.

Reference is now made to Fig. 2A and 2B. Figure 2A is a schematic illustration of a ballistic aircraft, generally referenced 120, constructed and operative in accordance with a preferred embodiment of the present invention. Figure 2B is a schematic illustration of a steering system, generally referenced 100, for mounting on the ballistic aircraft 120 of Figure 2A, constructed and operative in accordance with a preferred embodiment of the present invention.

Ballistic aircraft 120 includes a head 122, a body 126, a power nozzle 132, a rear end 130, an aft section 128, a stationary fin 102D and a groove 134. Groove 134 is only partially circular and ends with a wall 136 on side which is visible in the drawing and another respective wall, in the non visible side, which together define an arc, to which steering system 100 is to be mounted. With reference to Figure 2B, steering system 100 is shaped as a half circular arc, which includes three fins, referenced 102A, 102B and 102C.

Reference is further made to Figure 3, which is an illustration in detail of steering system 100 of Figure 2B. In accordance with the embodiment presented in Figure 3, fins 102A and 102C are dynamic, while fins 102B and 102D (Figure 2B) are stationary. As illustrated in Figure 3, fins 102A, 102B, 102C and 102D intersect a plane which is normal to a longitudinal axis of ballistic aircraft 120. Fin 102D is fixed to section 106 of ballistic aircraft 120 (Figure 2A). Steering system 100 is attached to section 106 of ballistic aircraft 120, by at least two fasteners 115 and 116. Section 106 is provided with a counterbore and a through hole both in a mutual axis, for each of the fasteners 115 and 116, whereby it is possible to insert each fastener in the counterbore and the through hole of section 106, and attach the fastener to a hole in steering system 100. Thus, steering system 100 is attached to section 106, and furthermore access is provided to fasteners 115 and 116, external to section 106. As can be seen in Figure 3, fin 102C is connected to steering system 100 housing by means of a shaft, generally referenced 104, which enables rotation thereof, with respect to an axis, defined by the shaft 104. It is noted that fin 102A is also connected to the steering system 100 housing, by means of a shaft (not shown) and hence is dynamic in a similar manner. Fin 102B is firmly fixed to housing of the steering system 100. It is noted that this fin can be firmly attached in a plurality of ways, such as welding and the like.

System 100 further includes avionics units, generally referenced 110, 112 and 114, located within the housing thereof. Avionics units 110, 112, and 114 each contain a controller and an actuator therein (not shown) to actuate and control the movement of dynamic fins 102A and 102C. These units thus control the position of each of the dynamic fins 102A and 102C, and hence, the direction of the ballistic aircraft, along its flight path. It is noted that since not all of the fins are dynamic, then the amount of fin actuators is reduced and hence, there is room for the avionics units within the housing of the steering system. Accordingly, locating the avionics units within the housing of the steering system simplifies the overall structure and eliminates the need for connecting the steering system with external avionics units, as in the prior art. It is noted that this aspect of the invention has a great implication over cost effectiveness as well as over reliability of the system of the invention, with respect to the prior art.

Reference is now made to Figures 4A, 4B and 5. Figure 4A is an illustration of system 100, at one mode of steering, operative in accordance with another preferred embodiment of the present invention. Figure 4B is an illustration of system 100, at another mode of steering, operative in accordance with another preferred embodiment of the present invention. Figure 5 is a schematic illustration of a method for operating the steering system, according to a further preferred embodiment of the present invention.

At every given moment, the maneuver required of the steering system is generally at a single direction, which is a vector summation of the desired maneuver in the pitch axis and the desired maneuver in the yaw axis. With reference to Figure 4A, vector 140 represents the pitch direction and vector 142 represents the yaw direction. Vector 144 represents the summation of both of vectors 140 and 142. In the present example, the steering system 100 is positioned where vector 144 is perpendicular to the joined axis of dynamic fins 102A and 102C. At this position, a rotation of both of these fins in the same direction (shown in Figure 4A), results in maneuvering the steering system and hence the ballistic object attached thereto, in the direction of vector 144.

With reference to Figure 4B, the steering system 100 can be directed to any selected position, with respect to the pitch and yaw axes. According to this aspect of the invention, the dynamic fins 102A and 102C are positioned at opposite directions, with respect to their mutual rotation axis. As can be seen from the drawing, fin 102A faces towards fin 102D, while fin 102C faces towards fin 102B. At this position, the steering system causes the attached ballistic object to roll. The roll direction can be set clockwise or counterclockwise. The angle of roll can be determined at any value. In accordance with the example, set forth in Figures 4A and 4B, the angle of roll is set to the difference between the current angle of fin 102B and direction of the summation vector 144. Accordingly, at any given moment of the flight of the ballistic object, the steering system can be rolled to a selected position and directed therefrom in the desired pitch and yaw directions.

This method of operation determines a slow rolling of the ballistic object. Hence, errors in the plane, which is perpendicular to the fins 102A and 102C can be accumulated by the navigation units within the steering system, and corrected thereby, as the ballistic object progresses along its course of flight.

With reference to Figure 5, the present invention provides a novel method for operating the steering system of the invention. According to this method, the steering system is constantly rolled into desired position and then maneuvered with respect to desired pitch and yaw directions. It is noted that according to a more advanced aspect of the invention, all of these operations are performed with some degree of overlapping. The initial step of the method of the invention includes detecting the current orientation of the ballistic object (step 200). Since the steering system 100 is firmly attached to the ballistic object 120, then this can be accomplished by determining the orientation of the steering system 100. In step 202, the system 100 determines a desired orientation for the ballistic object.

The desired orientation together with the current orientation, are then used to determine a required change in the pitch axis and the yaw axis (step 204).

The values of the changes in the pitch axis and the yaw axis are then used to determine a respective summation vect Mor (step 206). In step 208, the system 100 determines a roll direction from the current orientation of the steering system, towards the direction of the summation vector, where the plane defined by the dynamic fins is perpendicular to the direction of the summation vector 144 (see Figure 4A).

In step 210, the dynamic fins 102A and 102C are directed so as to roll the steering system, in the roll direction to the required position.

In step 212, the dynamic fins are directed so as to maneuver the steering system 100 according to the summation vector.

It is noted that steps 210 and 212 can be combined by directing each of the dynamic fins in a different and independent manner. For example, directing fin 102A to 5° with respect to the general axis of the ballistic object and directing fin 102C to -3° with respect to that same axis, yields a 3° equivalent roll of the ballistic object and a minor maneuver of 2° in an intermediate direction, with respect to the orientation of the steering system.

According to a further aspect of the invention, the control loop command can determined as follows: where SRC denotes the combined roll steering command, φ denotes a roll angle, KR1 and KR2 denote control loop gain values, p denotes the measured derivative of the roll value with respect to the axis of the ballistic object, ()c denotes a command (e.g., φ0 denotes a command roll angle), ()M denotes a measurement (e.g., φΜ denotes a measured roll angle) and ε( ) denotes an error.

It is noted that 5R denotes the combined roll angle and is determined according to the differential position of each of the dynamic fins J, (fin 102A) and δ^ (fin 102C), where δΗ = δ^ - δ .

Reference is further made to Figure 4C, which is a schematic view from the rear of the steering system 100 of Figure 2B, with the major steering parameters denoted thereon. The value of the acceleration commands in the two inertial planes, pitch and yaw. The required roll angle φ( is determined with respect to a vector combination of the two accelerations (apitCh and ayaw). The required maneuver ac in the current roll angle φ is determined from the acceleration command whose value is equal to the projection of the required maneuver, over the intermediate steering plane: a^VAL ≡ a^'- - COS(e≠) where ε denotes an error in the angular state of the object.

The combined steering command is determined from afWAL , as follows: hence, two separate commands δ . and δ . can be determined for each of the respective dynamic fins 102A and 102C, as follows: <V ~ 2 . 2 It is noted that the roll angle sr \s hence given by sr = £, - 3 and the combined pitch and yaw angle δ,,&ϊ is hence given by δ,,&γ = , +<S3.

Reference is now made to Figure 6, which is a schematic illustration of the avionics units, of system 100, constructed and operative in accordance with another preferred embodiment of the present invention.

Unit 110 includes a position detection unit 180, a guiding calculator 170, a control loop unit 172, a steering controller 174 and two actuators 176 and 178. The guiding calculator 170 is connected to the position detection unit 180 and to the control loop unit 172. The steering controller 174 is connected to the control loop unit 172 and to the actuators 176 and 178.

The position detection unit can include a plurality of types of position and orientation units, such as a global positioning system (GPS), acceleration detectors for each of the relevant axes (roll, pitch and yaw), an altitude detector, targeting detection devices, and the like.

The actuators can include any of a plurality of electromechanical devices, which are designed to change the position of a fin, connected thereto.

The position detection unit 180 detects a plurality of position and orientation parameters, with respect to the ballistic object and the environment in which it exists. These parameters can include the position x , the velocity x , time (denoted by the letter t), pitch, roll and yaw values and derivatives, and the like.

The position detection unit provides all of the detected and derived parameters (generally designated (ί, χ, ^,φ,ψ,θ, ρ, , ι·) ) to the guiding calculator 170. The guiding calculator determines an initial acceleration command ac and an initial roll angle command (. and provides these values to the control loop unit 172. The control loop unit 172 determines angle commands δκ: and 5M , and provides them to the steering controller 174. The steering controller determines steering commands sH. for the first actuator 176 and δ for the second actuator 178. The actuators 176 and 178 rotate the respective fins (102A and 102C) to the appropriate angles and provide respective feedback angles δ and δ therefrom. Hence, unit 170 constantly navigates the steering system 100 and the attached ballistic object towards its target, with respect to the intermediate physical parameters thereof.

Reference is now made to Figures 7A and 7B. Figure 7A is an illustration in perspective of a steering system, generally referenced 300, constructed and operative in accordance with another preferred embodiment of the present invention. Figure 7B is a view from the rear of system 300 of Figure 7A. System 300 includes two static fins 302B and 302D and two dynamic steering fins 302A, and 302C. Fins 302A, and 302C are mounted on shafts 304A (not shown) and 304C, respectively. It is noted that the same principals of the invention, which were discussed herein above, in conjunction with steering system 100, apply, with minor changes to system 300. Again, the structure of the steering system 300 simplifies the assembly stage, in which the steering system is to be added to a ballistic missile or bomb.

In accordance with a further aspect of the invention, there is provided yet a different system, which includes two dynamic fins and a single static one. Reference is now made to Figure 8, which is a schematic illustration of a steering system, generally referenced 350, constructed and operative in accordance with a further preferred embodiment of the present invention.

System 350 includes three fins 352A, 352B and 352C. Fin 352C is firmly fixed to the body of the ballistic object onto which the system 350 is attached. Fins 352A and 352B are connected to the housing of the system 350 by means of respective shafts 354A and 354B. Fins 352A, 352B and 352C are generally located at equal angles (120°) from each other, although it is noted that other arrangements are also possible.

According to another aspect of the invention, there is provided a method for extending the overall range of flight of a ballistic object. The method is characterized by lifting the ballistic object at some stage during descend path, so that it commences a new semi-launch at a height above the ground, and below the previous maximum.

Reference is now made to Figure 9, which is an illustration of side view of a flight path of a missile, according to another preferred embodiment of the present invention. In Figure 9, missile 402 is a conventional missile, and missile 404 is a missile constructed according to a preferred embodiment of the present invention. Missile 402 is launched from a site 406 towards a target 410, both located on a ground 420. Missile 404 is launched from site 406 towards a target 414, located on ground 420. Following the launch from site 406, missile 402 moves in a trajectory 411 up to a maximal point 409. After reaching the maximal point 409, missile 402 follows the trajectory 415 down toward target 410.

Powered flight path of missile 402, is from site 406 up to a point 408 above. From point 408, missile 402 follows an un-powered maneuvered path, to hit target 410 on ground 420. Alternatively, the entire flight path of missile 402 consisting of trajectories 411 and 415, no part of the flight path, or any combination of parts thereof may be powered or un-powered. The maximum range of missile 402 is a distance 416, being the shortest distance between points 406 and 410.

At a turning point in trajectory 415, generally referred to 412, missile 404 changes its normal trajectory 415, and instead, follows a path 422, which resembles a new launch at a point 417. It may be appreciated by those skilled in the art, that path 422 is made possible by rotating the dynamic fins 102A and 102B. Thence, missile 404 hits target 414, which is, located a distance 418 from launch site 406, greater than distance 416. The change of path might be caused, by utilizing the novel steering system described in the present invention.

Reference is now made to Figure 10, which is an illustration of side view of a flight path of a missile, according to another preferred embodiment of the present invention. Missile 434 is launched from site 436, and the flight path 437 is changed during a plurality of descend paths 447 and 453, at turning points 448 and 452. For example, missile 434 diverges from its normal trajectory 449 toward a maximal point 451 , by rotating the dynamic fins 102A and 102B, as described with respect to Figure 9. At a later stage in its flight path, missile 434 again diverges from its second normal trajectory 455, toward another maximal point 457, and then it hits a final target 454 located a distance 456 from launch site 436. It may be appreciated that distance 456 is greater than distance 450, and greater than distance 446, the maximum range of a conventional missile (not shown).

It will be appreciated by persons skilled in the art that the present invention is not limited to what has been particularly shown and described here in above. Rather the scope of the present invention is defined only by the claims which follow. 001191 -IL 134188/2

Claims (32)

1. Apparatus for controlling a flight path of a flying object, the object including at least one static fin, the apparatus comprising: a housing; and at least two dynamic fins, rotatably connected to said housing, wherein said at least one static fin and said at least two dynamic fins intersect a plane, said plane being normal to a longitudinal axis of said flying object.

2. The apparatus according to claim 1 , wherein said flying object is a self-propelled object.

3. The apparatus according to claim 1 , wherein said flying object is an un-propelled object.

4. The apparatus according to claim 1 , wherein said housing is fastened to said flying object, within a predetermined groove on said flying object.

5. The apparatus according to claim 1 , wherein said housing contains at least one avionics unit therein, for actuating said at least two dynamic fins. 001191 -IL 134188,

6. The apparatus according to claim 5, wherein each of said at least one avionics units controls actuation of said at least two dynamic fins.

Method for controlling a flight path of a flying object, said flying object comprising at least two dynamic fins7 and at least one static fin, said at least two dynamic fins and said at least one static fin intersecting a plane, said plane being normal to a longitudinal axis of said flying object, the method comprising the steps of: activating at least two of said dynamic fins such that said flying object substantially rolls about a longitudinal axis of said flying object; and activating at least two of said dynamic fins, thereby directing said flying object in a predetermined direction.

8. Method for controlling a flight path of a flying object, said flying object comprising at least two dynamic fins7 and at least one static fin, said at least two dynamic fins and said at least one static fin intersecting a plane, said plane being normal to a longitudinal axis of said flying object, the method comprising the steps of: detecting the orientation of said flying object in said flight path; determining a change in said orientation; and rotating said at least two dynamic fins according to said change.

9. The method according to claim 8, further comprising the step of determining a desired orientation. 001191-IL

10. The method according to claim 8, further comprising the step of determining a required orientation change, according to said desired orientation and said orientation.

11. The method according to claim 10, wherein said required orientation change includes a pitch change.

12. The method according to claim 10, wherein said required orientation change includes a yaw change.

13. The method according to claim 11 , wherein said required orientation change further includes a yaw change.

14. The method according to claim 12, wherein said required orientation change further includes a pitch change.

15. The method according to claim 10, further comprising the step of determining a summation vector, said summation vector being the vectorial sum of a pitch axis and a yaw axis determined according to said required orientation change.

16. The method according to claim 15, further comprising the step of determining a direction of roll, according to said summation vector. 001191 -IL 134188/2

17. The method according to claim 15, further comprising the step of determining a value of roll, according to said summation vector.

18. The method according to claim 15, further comprising the step of determining an angle in direction of said summation vector.

19. The method according to either of claims 8, 16, 17, and 18, further comprising the step of rolling said flying object according to said direction of roll, said value of roll, and said angle, by rotating said at least two dynamic fins.

20. The method according to either of claims 8, 15, and 19, further comprising the step of directing said flying object in direction of said summation vector, by rotating said at least two dynamic fins.

21. Flying object comprising: at least one static fin; a housing; and at least two dynamic fins, rotatably connected to said housing, wherein said at least one static fin and said at least two dynamic fins intersect a plane, said plane being normal to a longitudinal axis of said flying object. 001191 -IL 134188/2

22. The flying object according to claim 21 , wherein said flying object is a self-propelled object.

23. The flying object according to claim 21 , wherein said flying object is an un-propelled object.

24. The flying object according to claim 21 , wherein said housing is fastened to said flying object, within a predetermined groove on said flying object.

25. The flying object according to claim 21 , wherein said housing contains at least one avionics unit therein, for actuating said at least two dynamic fins.

26. The flying object according to claim 25, wherein each of said at least one avionics units controls actuation of said at least two dynamic fins.

27. Apparatus for controlling a flight path of a flying object according to any of claims 1 - 6 substantially as described hereinabove.

28. Apparatus for controlling a flight path of a flying object according to any of claims 1 - 6 substantially as illustrated in any of the drawings. 001191-IL 134188/2

Method according to any of claims 7 - 20 substantially as described hereinabove.

30. Method for controlling a flight path of a flying object according to any of claims 7 - 20 substantially as illustrated in any of the drawings.

31 Flying object according to any of claims 21 - 26 substantially as described hereinabove.

32. Flying object according to any of claims 21 - 26 substantially as illustrated in any of the drawings. For the applicant BOROCHOV, KORAKH, ELIEZRI & CO., Advocates & Patent Attorneys Eliav Korakh, Advocate & Patent Attorney