US3210025A - Empennage construction for a space missile - Google Patents

Empennage construction for a space missile Download PDF

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US3210025A
US3210025A US129776A US12977661A US3210025A US 3210025 A US3210025 A US 3210025A US 129776 A US129776 A US 129776A US 12977661 A US12977661 A US 12977661A US 3210025 A US3210025 A US 3210025A
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case
missile
arm
movement
position
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US129776A
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George F Lubben
Edward J Sampson
Jr Donald W Robinson
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KAMAN AIRCRAFT CORP
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KAMAN AIRCRAFT CORP
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    • BPERFORMING OPERATIONS; TRANSPORTING
    • B64AIRCRAFT; AVIATION; COSMONAUTICS
    • B64GCOSMONAUTICS; VEHICLES OR EQUIPMENT THEREFOR
    • B64G1/00Cosmonautic vehicles
    • B64G1/22Parts of, or equipment specially adapted for fitting in or to, cosmonautic vehicles
    • B64G1/62Systems for re-entry into the earth's atmosphere; Retarding or landing devices
    • BPERFORMING OPERATIONS; TRANSPORTING
    • B64AIRCRAFT; AVIATION; COSMONAUTICS
    • B64CAEROPLANES; HELICOPTERS
    • B64C29/00Aircraft capable of landing or taking-off vertically
    • B64C29/02Aircraft capable of landing or taking-off vertically having its flight directional axis vertical when grounded

Description

Oct. 5, 1965 G. F. LUBBEN ETAL EMPENNAGE CONSTRUCTION FOR A SPACE MISSILE Filed Aug. 7, 1961 4 Sheets-Sheet 1 ATTORNEYS EMPENNAGE CONSTRUCTION FOR A SPACE MISSILE Filed Aug. 7, 1961 O 5, 9 G. F. LUBBEN ETAL 4 Sheets-Sheet 2 m wt a F .aIM-" -I 1965 G. F. LUBBEN E' l'AL EMPENNAGE CONSTRUCTION FOR A SPACE MISSILE Aug. 7. 1961 R 4 Sheets-Sheet 5 Oct. 5, 1965 G. F. LUBBEN ETAL 3,210,025

EMPENNAGE CONSTRUCTION FOR A SPACE MISSILE Filed Au 7, 1961 72 74 4 Sheets-Sheet 4 United States Patent corporation of Connecticut Filed Aug. 7, 1961, Ser. No. 129,776 14 Claims. (Cl. 244-1) This invention relates to an empennage construction having one or more aerodynamic surfaces for stabilizing the movement of a space missile during its flight through the atmosphere, and deals more particularly with an empennage construction for a space missile adapted to make a relatively soft landing on the earth by following a glide path during the terminal portion of its descent.

The empennage construction of this invention finds particular utility in conjunction with a space missile of the type having a rotor for retarding its descent and which rotor is adjustable to vary the flight path of the missile as it approaches the earth. Such a missile may be released in mid-air or outer space in a variety of ways, as from an aircraft or other space missile, and the rotor autorotates as air flows past it in such a manner as to apply a retarding force to the missile. By means of a control unit in the missile actuated by guidance signals transmitted thereto from the ground, or by signals from a self-contained and preprogrammed guidance system, the rotor may be adjusted to cause the missile to follow a controlled glide path as it approaches the earth. Throughout the glide path the missile travels in a generally sidewise direction, as contrasted to the earlier portions of its descent during which it travels in a generally lengthwise direction.

Because the empennage construction of this invention is particularly useful in a missile such as aforesaid it has been shown and described herein, by way of example, as included in such a missile. It is to be understood, however, that by doing so there is no intention to limit the invention, at least in its broader aspects, to a missile hav ing a retarding rotor, but that certain features of the invention may be applied as well to other types of missiles.

The term space missile as used herein is used to refer both to devices which travel in outer space and re-enter the atmosphere on landing and also to strictly aerial devices which are intended for flight only in the atmosphere.

One object of this invention is to provide a means for stabilizing the movement of a space missile about its central axis as the missile travels either in a lengthwise direction parallel to its central axis or in a sidewise direction generally perpendicular to the central axis.

A more particular object of this invention is to provide an empennage construction for a space missile which empennage construction is adapted to prevent spin of the missile as it travels in a path parallel to its longitudinal axis and is also adapted to provide directional stability for the missile as it travels in a glide path perpendicular to its longitudinal axis.

Another object of this invention is to provide an empennage construction adapted for use with a space missile for stabilizing its movement during its flight through the atmosphere and for serving as a turnover gear upon landing to hold the missile in an upright position.

Other objects and advantages of the invention will be apparent from the following description and the accompanying drawings.

The drawings show a preferred embodiment of the invention and such embodiment will be described, but it will be understood that various changes may be made from the construction disclosed, and that the drawings and description are not to be construed as defining or limiting the scope of the invention, the claims forming a part of this specification being relied upon for that purpose.

Of the drawings:

FIG. 1 is a schematic view illustrating a typical flight path of a space missile including an empennage construction embodying the present invention.

FIG. 2 is an elevational view of the space missile shown in FIG. 1.

FIG. 3 is an enlarged fragmentary elevational view of the space missile of FIG. 1 with the illustrated empennage arm being shown midway between its normal and its deployed positions.

FIG. 4 is an enlarged plan view of the empennage arm which carries a fixed fin or stabilizer.

FIG. 5 is an enlarged transverse sectional view taken on the line 55 of FIG. 4.

FIG. 5a is an enlarged fragmentary sectional view taken through the solenoid operated pin of FIG. 5.

FIG. 6 is an enlarged sectional view taken on the line 66 of FIG. 4, with the view being drawn on a scale larger than that of FIG. 4 but smaller than that of FIG. 5.

FIG. 7 is an enlarged horizontal sectional view taken on the line 77 of FIG. 2.

FIG. 8 is an enlarged fragmentary plan view of the rear end portion of one of the empennage arms which carries a pivoting fin or stabilizer.

FIG. 9 is a transverse sectional view taken on the line 9-9 of FIG. 8.

General organization and flight pattern of space missileFlGS. 1 and 2 The flight pattern or profile of a space missile embodying the present invention is shown in FIG. 1, while FIG. 2 shows the general organization or structure of the missile. Referring to these figures, it will be noted that the missile, indicated generally at It), includes a generally elongated case 12 having a hexagonal shape in cross section. The nose of the case is rounded and carries an antenna 14 which may be used to receive radio command signals and/or as part of a radar altimeter. It is to be understood, however, that the shape of the case 12 may vary widely without departing from the invention and that it need not necessarily be an elongated structure, although such configuration is preferred.

For retarding the descent of the missile 10, the missile case 12 has a rotor assembly, indicated generally at 16, attached to its rear end. The rotor assembly 16 may take various different forms, and may be generally similar to that shown in the patent to Robinson et al., No. 3,017,147, to which patent reference is made for a more detailed description. For the purpose of discussion, however, and in the illustrated case, the rotor assembly 16 includes two rotor blades 18, 18 supported'for rotation about a central shaft 20. Before release from the aircraft or other vehicle which carries the missile into the atmosphere or outer space, the blades 18, 18 are folded inwardly, as shown by the broken lines of FIG. 2, so as to extend generally longitudinally or parallel to the axis 'of the case 12 and the shaft 20. After the missile is released, other mechanism, surrounding the shaft 20 and indicated generally at 22, operates to release the blades 13, 18 from the folded position shown by the broken lines of FIG. 2 and to cause the blades to pivot about their forward ends for some distance, thereby moving the rear ends of the blades outwardly and away from the axis of the shaft 20. This operation is referred to as pre-coning the blades and in performing this operation the mechanism 22 may additionally impart some initial rotation to the blades.

After the blades 18, 18 are pre-coned, the air moving past the blades exerts aerodynamic forces thereon to cause the same to autorotate about the shaft 20. As the speed of autorotation increases the centrifugal forces acting on the blades cause the latter to pivot further about their forward ends and to move to the extended or generally transverse positions shown by the solid lines of FIG. 2. When extended, it is important that the two blades 18, 18 have uniform coning angles with respect to the axis of the shaft 20, and for this purpose the rotor mechanism includes an equalizing apparatus comprising a collar 24 loosely mounted on the shaft 20 and connected respectively with the two blades 18, 18 by two links 26, 26.

As the blades 18, 18 reach their extended or operative positions, the autorotation thereof causes a braking or retarding force to be applied to the missile to slow its vertical descent. The magnitude of the retarding force depends on the speed of the rotor rotation and the negative pitch angle of the rotor blades, and preferably, but not necessarily, the mechanism 22 includes means for controlling the pitch angles of the rotor blades in order to obtain an optimum rotor speed and an optimum retarding force. Also, the rotor preferably includes means whereby the pitches of the rotor blades 18, 18 may be changed from negative to positive valves to provide for collective flare on landing 'on command from control devices located within the instrument case 12.

In the illustrated case, the rotor 16 is connected to the rear end of the case 12 by means of a universal joint 28 which allows the rotor shaft'20 to be tilted in any direction relative to the central or longitudinal axis of the case 12. This tilting movement of the rotor is controlled by two control rods 30, 30 operated by servo motors or the like within the case 12. Tilting the rotor by tilting the axis of the shaft 20 relative to the axis of the case 12 is equivalent to inducing a cyclic pitch change in the rotor blades 18, 18 and accordingly may be used to control the direction of flight of the missile 10 in a manner similar to the use of cyclic pitch changes to control the flight of a helicopter.

FIG. 1 shows the flight path of a missile 10 which may be assumed to be released from an airplane. In this case, therefore, the missile 10 travels initially in a generally horizontal path as shown at a in FIG. 1. At this time the blades are in their folded positions. Shortly after the release from the airplane the rotors are deployed by the operation of the pre-coning mechanism followed by further coning as a result of centrifugal forces. At 12 in FIG. 1 the rotor blades are shown in the process of being deployed and are midway between their folded and their fully deployed or operative positions. From FIG. 1 it will be understood that upon movement of the blades to their operative positions, the attitude of the case quickly shifts from the generally horizontal position shown at a and b to an approximately vertical position as shown at 0.

So long as no tilting movement is applied to the rotor 16, the missile 10 follows a generally lengthwise flight path. That is, it travels in a direction generally parallel to its longitudinal axis. In a typical flight path this lengthwise movement of the missile is maintained until the missile reaches a predetermined low altitude. In other words, after the rotor is deployed the missile usually drops in a generally vertical path towards the ground until it gains a certain low altitude. At in FIG. 1 the missile is shown at the end of this generally vertical descent.

After the missile reaches this stage of its flight, preprogrammed controls within the case 10, or controls responsive to radio signals transmitted from a ground control station 32, are brought into play to tilt the axis of the rotor shaft 20 relative to the case 12. This tilting of the rotor causes the missile to leave its generally vertical path and to partake of a gliding movement in which the case 12 moves in a generally sidewise direction while nevertheless retaining a generally vertical attitude. The appearance of the missile 10 while traveling in such a glide path is shown at d in FIG. 1.

During the lengthwise movement of the missile 1%, or during the portion of the flight path from a to c in FIG. 1, the autorotation of the rotor 15 imparts a torque on the case 12 tending to rotate the latter in the same direction as that of the blade rotation. That is, the case 12 tends to turn or spin about its longitudinal or central axis. A similar torque is also imposed on the case during its sidewise movement along the glide path. Rotation of the case during the gliding motion of the case is especially undesirable since in order for the controls within the case to operate properly it is necessary that the same side of the case always be pointed forwardly. Due to the difference in the direction of movement of the case, the stabilizing means which would ordinarily be used to stabilize the case movement during lengthwise movement is incapable of stabilizing its movement during the sideways movement which occurs throughout the glide path. One of the features of this invention is therefore an empennage construction which includes a stabilizing means and is movable between a first position at which the stabilizing means acts to properly stabilize the missile during lengthwise movement and a second position at which the stabilizing means serves to stabilize the case during the sideways movement occurring within the glide path.

if he empennage construction is shown generally in FIG. 2 and in more detail in FIGS. 3 to 9. Before considering the structure of the empennage, however, it should be noted that one of the features of the rotor equipped missile 10 is the ability of the missile to he landed softly so as not to damage the contents of the case 12. FIG. 1 shows the flight path which may be followed to obtain such a soft landing. As noted previously, after the missile reaches the terminal portion of its vertical descent it is started in a glide path by a tilting motion of the rotor 16. The missile case 12 thereafter travels in a generally sidewise direction while retaining its vertical attitude. While traversing this glide path the missile may be guided to a selected landing area by signals transmitted from the ground station 32. Throughout this glide path the rotor 16, in order to maintain autorotation and the forward travel of the case 12, is tilted generally forwardly from the axis of the case, as will be apparent from the showing at d in FIG. 1. Also, during this glide path, it will be apparent that the missile will develop a substantial forward ground speed as well as a vertical speed. In order to obtain a soft landing both the forward ground speed and the vertical speed should be reduced as far as possible towards zero at the moment of impact.

To achieve this desirable reduction in the forward and vertical speeds, the missile 10 as it approaches the ground is first controlled to cause the rotor 16 to tilt rearwardly with respect to the axis of the case 12. This is referred to as cyclic flare, as shown at e, and causes a reduction in the forward or ground speed of the missile. As the ground speed approaches zero as a result of the cyclic flare, the missile tends to drop in a purely vertical direction and at this time the inertia of the moving blades may be used to apply a sudden large lifting force to the missile case by collectively flaring the rotor blades. That is, the pitches of the two blades 18, 18 are suddenly changed from their normal negative pitch positions, used to achieve autorotation, to positive pitch positions. Thus the missile case 12 may he landed softly and in a vertical position as shown at f in FIG. 1.

Once the missile is landed in a vertical position as shown in FIG. 1 it is desirable that the case be maintained in the vertical position and prevented from falling over on its side, since such a fall could damage the contents of the missile case. In this respect, another feature of the empennage construction of this invention is that it serves as a turnover gear to hold the missile case 12 in its vertical position after landing.

Referring to FIG. 2, the empennage construction of this invention, as applied to the illustrated missile 10, comprises three elongated arms 34, 34 pivotally connected at their forward ends to the missile case 12, one of the arms 34 being hidden from the view in FIG. 2. The arms 34, 34 are pivoted for movement about axes arranged generally transverse to the longitudinal axis of the case 12 and are movable between first positions at which the arms are arranged adjacent the side of the case with their longitudinal axes generally parallel to the central or longitudinal axis of the case 12, as shown by the broken lines of FIG. 2, and second positions at which the arms extend outwardly and generally transverse to the case 12, as

shown by the solid lines of FIG. 2. Attached to the free or rear end of each arm 34 is a stabilizer in the form of a fin 36. When the arms 34, 34 are in their first or normal positions, the associated fins 36, 36 extend outwardly and longitudinally of the case 12 and act in a conventional manner to stabilize movement of the case 12 about its vertical axis during its lengthwise movement.

When the three arms are moved to their second or deployed positions, one of the fins 36, shown at the right in FIG. 2, remains fixed relative to its arm 34 and serves in a fashion similar to the vertical stabilizer of an air plane for stabilizing movement about the vertical axis of the case 12 as the case travels in a sidewise direction. The other two fins, 36, 36 are allowed to pivot relative to their associated arms 34, 34 so as to weathervane to reduce drag and permit the fixed fin 36 to take over the entire stabilization function.

Once the arms 34, 34 are moved to their extended or deployed positions they are held locked in that position so that after the missile lands the arms act as the legs of a turnover gear for holding the missile case in a substantially upright position.

The details of the empennage assembly shown in FIG. 2 will be described in more detail hereinafter in conjunction with FIGS. 3 to 9. It should be noted, however, that the details of construction may be varied widely without departing from the invention. It should be particularly noted that in cases where it is unnecessary to provide for a turnover gear it may be desirable to provide only one pivoting arm similar to the arm 34 shown at the right in FIG. 2. Also, regardless of whether one or more pivoting arms are provided it may be desirable to connect the stabilizing means, such as the three fins 36, 36 to one of the arms 34 so that this arm carries with it the entire stabilizing means when it is moved to its deployed position.

Detailed description of empelmage construction FIGS. 3 to 9 The means for moving the arms 34, 34 between their normal and their deployed positions is shown best in FIGS. 3 and 4, which show the one arm having the fixed fin 36 associated therewith. Except for the fixed fin, however, this arm and its associated structure is similar to the other two arms. Referring to these figures, it will be noted that the arm 34 is pivotally connected at its forlward end to the case 12 by means of a bracket 38 for movement about a transverse axis. Also connected to the side of the case 12 is a longitudinally extending track 40, which track cooperates with a slide 42. The track and slide may take various different forms, but as shown Between the slide 42 and the arm 34 is a strut 50 which is pivotally connected at one end to the forward end of the slide 42 and pivotally connected at its other end to the arm 34 at a point intermediate the ends of the arm. It will therefore be obvious from FIG. 3, that sliding movement of the slide 42 along the track 40 will cause the strut 50 to move the arm 34 about its pivot axis. As viewed in FIG. 3, downward movement of the slide 42 will cause the arm 34 to be moved toward its extended position. Preferably, as shown in FIG. 6, the arm 34 is of a circular cross section and the strut 50 of a channel shaped cross section so that the arm 34 may nest within the strut 50 while the arm 34 is in its normal position.

Associated with the track 40 is a stop 52 and a latch 54. The latch 54 is of the type which permits the slide 42 to move downwardly past .the latch but prevents the slide from returning in the opposite direction. The stop 52 is so located that when engaged by the slide 42 the arm 34 will be held in approximately the deployed or transverse position shown by the solid lines of FIG. 2. The latch 54 at the same time prevents the slide 42 from travelling back along the track 40 so that once the arm 34 is moved to its deployed position it is held in such position by the latch 54 and the stop 52.

Means, hereinafter described, are provided for initially holding the arm 34 in its first or normal position adjacent the side of the missile case 12, together with means for biasing the arm from its normal to its deployed position. In the illustrated case the biasing means consists of an elastic shock cord 56. The shock cord 56 is stretched between a bracket 58, fastened to the side of the case 12, and the slide 42 so as to pull the slide 42 downward when the holding means for the arm 34 is released.

As mentioned previously, FIGS. 3 and 4 show the arm 34 which appears at the right in FIG. 2 and which includes the fixed fin 36. This same fin appears at the upper right-hand portion of FIG. 7; and, from FIGS. 3, 4 and 7 it will be noted that the fin consists of a relatively thin and flat plate which is fixedly secured to the arm 34 by means of two retaining members 60, 60, with each retaining member being riveted to the arm 34 and bolted to the fin 36 as shown best in FIG. 7. This fin 36 is so positioned on its associated arm 34 as to extend outwardly and longitudinally of the case 12 and to be disposed in a plane passing through the longitudinal axis of the case when the arm 34 is in its normal position.

The structure of the other .two fins 36, 36, which fins are allowed to pivot on their associated arms 34, 34 when the arms are in their deployed positions, is shown in FIGS. 8 and 9. Each of these fins is, similarly to the first fin 36, comprised of a relatively thin, flat plate. Associated with each pivoting fin is a tubular retaining member 62 which loosely surrounds the associated cylindrical arm 34. Welded to the tubular member 62 and forming a part thereof are two outwardly extending spaced plates 64, 64, the fin 36 being received between the two plates and bolted thereto. The tubular member 62 also has welded thereto gusset pieces 66, 66 which provide additional support for the plates 64, 64. The tubular member 62 is axially retained in place on the arm 34 by two collars 68, 68 fastened to the arm.

When the arm 34 of FIGS. 8 and 9 is in the normal position shown, the associated fin 36 is held fixed in an outwardly extending position relative to the case 12 by means of two pins 70, 78 carried by the associated strut 50. As shown best in FIG. 9, the pins 70, 70 extend through registering openings in the tubular member 62 and in the arm 34 to hold the tubular member and fin in a fixed angular position relative to the arm. When the arm 34 is moved from its normal position towards its deployed position the pins 70, 70 are removed from the openings thereby freeing the tubular member 62 and the fin 36 for pivoting or Weathervaning motion relative to the arm.

The three arms 34, 34 are held in their first or normal positions by releasable means which may be operated by a control unit within the case 12 to release the arms when the missile 10 reaches a predetermined altitude and begins its gliding movement. In the illustrated case this means comprises three solenoid devices, indicated at 72, 72, and best shown in FIG. 7. The solenoid devices 72, 72 are supported by brackets 74, 74 fastened to the end wall or plate 75 of the missile case. Extending outwardly from each bracket 74 is a pin 76 which passes through an opening formed in a plate 78 welded to the rear end of the associated arm 34, as best shown in FIGS. and Set. On the outer end of each pin 76 are two spring-biased balls 80, 80 which bear against a coneshaped seat 82 formed in the plate 78 and normally releasably holds the arm 34 in its first or normal position. Upon actuation of the solenoid, however, .the pin 76 is pulled inwardly toward the center of the case 12 by a force sufiicient to cause the balls 80, 80 to be camrned inwardly toward the center of the pin 76 by the seat 82, thereby allowing the pin to be freed from the plate 78.

Surrounding each pin 76 is a helical compression spring 84- which functions when the pin is pulled from the associated plate 7 8 to urge the plate and arm 34 outwardly to initiate its movement to the deployed position. The associated shock cord 56 thereafter moves the arm to the fully deployed position, in which position it is held by the associated stop 52 and latch 54.

The invention claimed is:

1. A space missile adapted for landing on the earth after a descending flight through the atmosphere, said missile comprising an elongated missile case having a central axis extending longitudinally thereof, a rotor attached to the rear end of said case for retarding the descent of said missile during its flight through the atmosphere and which rotor is selectively operable to cause said missile case to travel either in an endwise direction generally parallel to said central axis or in a side wise direction generally perpendicular to said central axis, an elongated arm having one end connected with said case for movement of said arm between a first position at which the longitudinal axis of said arm is arranged generally parallel with the central axis of said case and a second position whereat said arm extends outwardly from said case with its longitudinal axis generally perpendicular to said central axis, and a stabilizer connected to the free end of said arm for stabilizing movement of said case about said central axis when said case travels generally parallel to said central axis with said arm in said first position and for similarly stabilizing movement of said case about said central axis when said case travels generally perpendicular to said central axis with said arm in said second position.

2. A space missile as defined in claim 1 further characterized by means for releasably holding said arm in said first position, and means for moving said arm from said first position to said second position when said latter means is operated to release the same.

3. A space missile as defined in claim 2 further characterized by latch means for holding said arm in said second position and for preventing its return to said first position once moved to said second position.

4. A space missile adapted for landing on the earth after a descending flight through the atmosphere, said missile comprising an elongated missile case having a longitudinal axis, a rotor attached to the rear end of said case for retarding the descent of said missile during its flight through the atmosphere and which rotor is selectively operable to cause said missile case to travel either in an endwise direction generally parallel to said longitudinal axis or in a sidewise direction generally perpendicular to said axis, an elongated arm pivotally connected at its forward end to the side of said case adjacent the forward end of the latter for movement about a transverse axis between a normal position at which said arm extends rearwardly from its pivotal connection with said case and is disposed adjacent the side of said case in generally parallel relationship with the longitudinal axis thereof and a deployed position whereat said arm extends outwardly from said forward end of said case, and a stabilizer con nected to the rear end portion of said arm for stabilizing movement of said case about its longitudinal axis when said case travels in said endwise direction with said arm in said normal position and for similarly stabilizing movement of said case about its longitudinal axis when said case travels in said sidewise direction with said arm in said deployed position.

5. A space missile as defined in claim 4 further characterized by said stabilizer comprising a fin fixed to the rear end portion of said arm and arranged to extend outwardly from said arm and parallel to the longitudinal axis of said case when said arm is in said normal position adjacent the side of said case.

6. A space missile as defined in claim 4 further characterized by a longitudinally extending track connected with the side of said case, a slide carried by said track for movement along its length, and a strut pivotally connected at one end to said slide and at its other end to said arm at a point intermediate the ends of said arm so that said arm is moved by said strut from its normal to its deployed position as said slide is moved from a first to a second position along the length of said track.

7. A space missile as defined in claim 6 further characterized by means for biasing said slide toward said second position, means for releasably holding said arm in its normal position and said slide in its first position against the force of said biasing means, and latch means associated with said track for cooperating with said slide to hold said arm in its deployed position and said slide in its second position once said slide is moved to said second position by the action of said biasing means.

8. A space missile as defined in claim 7 further characterized by said means for biasing said slide comprising a shock cord fastened at one end of said slide and at its other end to said case and adapted to pull said slide from said first to said second position.

9. In a space missile the combination comprising a missile case having a central axis, a plurality of elongated arms angularly spaced about said central axis and each of which arms is connected with said case for movement between a normal position at which the longitudinal axis of said arm is arranged generally parallel with the central axis of said case and a deployed position whereat said arm extends outwardly from said case with its longitudinal axis generally perpendicular to said central axis, biasing means for urging said arms from their normal to their deployed positions, means for releasably holding said arms in their normal positions against the action of said biasing means, said arms being moved to said deployed position by said biasing means when said holding means is released, and latch means for holding said arms in their deployed positions once moved to said deployed positions by said biasing means so that said arms serve to hold said case in a generally upright position on the ground when said missile is landed with said central axis generally vertical and with said arms in their deployed positions.

10. The combination as defined in claim 9 further characterized by selectively operable means for causing said case to travel either in an endwise direction generally parallel to said central axis or in a sidewise direction generally perpendicular to said central axis, and a stabilizer fixed to one of said arms for stabilizing movement of said case about said central axis when said case travels in the atmosphere along a flight path generally parallel to said central axis with said arms in their normal positions and to similarly stabilize movement of said case about said central axis when said case travels along a flight path generally perpendicular to said central axis with said arms in their deployed positions.

11. A space missile comprising a missile case, a rotor connected with said case for retarding its descent during flight through the atmosphere and selectively operable to cause said case to travel in either a generally lengthwise direction with its central axis vertical or in a generally sidewise direction with its central axis vertical, and an empennage construction for stabilizing the movement of said case about said central axis, said empennage construction being movable relative to said case between a first position whereat it is eifective to stabilize said case in regard to movement about said central axis during generally lengthwise movement thereof and a second position whereat it is similarly efiective to stabilize said case during generally sidewise movement thereof.

12. A space missile as defined in claim 11 further characterized by said empennage including a plurality of arms which arms extend outwardly from said case at various angularly spaced positions about said central axis when said empennage is in said second position so as to support said missile case in a generally upright position after landing.

13. A space missile as defined in claim 11 further char acterized by said empennage construction including an elongated arm having one end connected with said case for movement of said arm between a first position at which said arm is disposed adjacent the side of said case in generally parallel relationship to the central axis thereof and a second position whereat said arm extends outwardly from said case, and a stabilizer fixed to the rear end portion of said arm.

14. An empennage construction for a space missile, said empennage construction comprising a missile case having a central axis, a plurality of elongated arms angularly spaced about said central axis and each of which arms is connected with said case for movement between a normal position at which the longitudinal axis of said arm is arranged generally parallel with the central axis of said case and a deployed position whereat said arm extends outwardly from said case with its longitudinal axis generally perpendicular to said central axis, means for releasably holding said arms in their normal positions, means for moving said arms from their normal to their deployed positions when said latter means is operated to release the same, latch means for holding said arms in their deployed positions once moved to said deployed positions so that said arms serve to hold said case in a generally upright position on the ground when said missile is landed with said central axis generally vertical and with said arms in their deployed positions, a plurality of fins each associated with a respective one of said arms, one of said fins being fixedly secured to its associated arm and the other of said fins being pivotally secured to their associated arms, and means for holding said other fins in a fixed position relative to their associated arms when said arms are in their normal positions so that all of said fins act collectively to stabilize movement of said case about said central axis when said case travels in a flight path generally parallel to its central axis, said latter means being operable to free said other fins for pivoting movement on their associated arms when said arms are moved to their deployed positions so that said fixed fin acts to stabilize movement of said case about its central axis while said other fins move in a weathervaning manner when said case travels along a flight path generally perpendicular to said central axis with said arms in their deployed positions.

References Cited by the Examiner UNITED STATES PATENTS 1,463,471 7/23 Kadel 244102 1,840,152 1/32 Buchanan 244-1719 FOREIGN PATENTS 1,005,158 12/51 France.

FERGUS S. MIDDLETON, Primary Examiner.

Claims (1)

11. A SPACE MISSILE COMPRISING A MISSILE CASE, A ROTOR CONNECTED WITH SAID CASE FOR RETARDING ITS DESCENT DURING FLIGNT THROUGH THE ATMOSPHERE AND SELECTIVELY OPERABLE TO CAUSE SAID CASE TO TRAVEL IN EITHER A GENERALLY LENGHTWISE DIRECTION WITH ITS CENTRAL AXIS VERTICAL OR IN A GENERALLY SIDEWISE DIRECTION WITH ITS CENTRAL AXIS VERTICAL, AND AN EMPENNAGE CONSTRUCTION FOR STABILIZING THE MOVEMENT OF SAID CASE ABOUT SAID CENTRAL AXIS, SAID EMPENNAGE CONSTRUCTION BEING MOVABLE RELATIVE TO SAID CASE BETWEEN A FIRST POSITION WHEREAT IT IS EFFECTIVE TO STABILIZE SAID CASE IN REGARD TO MOVEMENT ABOUT SAID CENTRAL AXIS DURING GENERALLY LENGHTWISE MOVEMENT THEREOF AND A SECOND POSITION WHEREAT IT IS SIMILARLY EFFICTIVE TO STABLIZE SAID CASE DURING GENERALLY SIDEWISE MOVEMENT THEREOF.
US129776A 1961-08-07 1961-08-07 Empennage construction for a space missile Expired - Lifetime US3210025A (en)

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Cited By (12)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3485466A (en) * 1966-12-27 1969-12-23 Richard H Prewitt Rotary wing device
US3623682A (en) * 1968-09-27 1971-11-30 Giraviens Dorand Rotary wing aircraft
US3807671A (en) * 1973-03-05 1974-04-30 Stencel Aero Eng Corp Escape and recovery system
EP0321721A2 (en) * 1987-12-22 1989-06-28 ERNO Raumfahrttechnik Gesellschaft mit beschränkter Haftung Device to assist in the landing of heavy space capsules
US5842665A (en) * 1996-09-09 1998-12-01 Hmx, Inc. Launch vehicle with engine mounted on a rotor
US5873549A (en) * 1996-09-25 1999-02-23 Mcdonnell Douglas Corporation Vehicle rotation and control mechanism
US20040099768A1 (en) * 2002-11-22 2004-05-27 Maryan Chak Aircraft, with means for at least reducing impact against the ground
US20100327107A1 (en) * 2009-02-24 2010-12-30 Blue Origin, Llc Bidirectional control surfaces for use with high speed vehicles, and associated systems and methods
US20110017872A1 (en) * 2009-06-15 2011-01-27 Blue Origin, Llc Sea landing of space launch vehicles and associated systems and methods
JP2012518575A (en) * 2009-02-24 2012-08-16 ブルー オリジン エルエルシー Launch vehicle with stationary and deployable deceleration surface and / or profile fuel tank and related systems and methods
WO2014177589A1 (en) * 2013-04-30 2014-11-06 Johannes Reiter Aircraft for vertical take-off and landing with hinged and bendable wings
US9896196B1 (en) * 2014-05-21 2018-02-20 Kaiser Enterprises, Llc Manned and unmanned aircraft

Citations (3)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US1463471A (en) * 1921-10-06 1923-07-31 George W Kadel Landing carriage for aeroplanes
US1840152A (en) * 1931-02-16 1932-01-05 William J Buchanan Flying machine
FR1005158A (en) * 1947-06-12 1952-04-07 Airplane at high speed

Patent Citations (3)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US1463471A (en) * 1921-10-06 1923-07-31 George W Kadel Landing carriage for aeroplanes
US1840152A (en) * 1931-02-16 1932-01-05 William J Buchanan Flying machine
FR1005158A (en) * 1947-06-12 1952-04-07 Airplane at high speed

Cited By (20)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3485466A (en) * 1966-12-27 1969-12-23 Richard H Prewitt Rotary wing device
US3623682A (en) * 1968-09-27 1971-11-30 Giraviens Dorand Rotary wing aircraft
US3807671A (en) * 1973-03-05 1974-04-30 Stencel Aero Eng Corp Escape and recovery system
EP0321721A2 (en) * 1987-12-22 1989-06-28 ERNO Raumfahrttechnik Gesellschaft mit beschränkter Haftung Device to assist in the landing of heavy space capsules
DE3743513A1 (en) * 1987-12-22 1989-07-13 Erno Raumfahrttechnik Gmbh Apparatus for landing aid of heavy space capsules
EP0321721A3 (en) * 1987-12-22 1990-07-04 ERNO Raumfahrttechnik Gesellschaft mit beschränkter Haftung Device to assist in the landing of heavy space capsules
US5842665A (en) * 1996-09-09 1998-12-01 Hmx, Inc. Launch vehicle with engine mounted on a rotor
US5873549A (en) * 1996-09-25 1999-02-23 Mcdonnell Douglas Corporation Vehicle rotation and control mechanism
US20040099768A1 (en) * 2002-11-22 2004-05-27 Maryan Chak Aircraft, with means for at least reducing impact against the ground
US8894016B2 (en) 2009-02-24 2014-11-25 Blue Origin, Llc Bidirectional control surfaces for use with high speed vehicles, and associated systems and methods
US20100327107A1 (en) * 2009-02-24 2010-12-30 Blue Origin, Llc Bidirectional control surfaces for use with high speed vehicles, and associated systems and methods
JP2012518575A (en) * 2009-02-24 2012-08-16 ブルー オリジン エルエルシー Launch vehicle with stationary and deployable deceleration surface and / or profile fuel tank and related systems and methods
US8991767B2 (en) 2009-02-24 2015-03-31 Blue Origin, Llc Control surfaces for use with high speed vehicles, and associated systems and methods
US8878111B2 (en) 2009-02-24 2014-11-04 Blue Origin, Llc Bidirectional control surfaces for use with high speed vehicles, and associated systems and methods
US8876059B2 (en) 2009-02-24 2014-11-04 Blue Origin, Llc Bidirectional control surfaces for use with high speed vehicles, and associated systems and methods
US9580191B2 (en) 2009-02-24 2017-02-28 Blue Origin, Llc Control surfaces for use with high speed vehicles, and associated systems and methods
US20110017872A1 (en) * 2009-06-15 2011-01-27 Blue Origin, Llc Sea landing of space launch vehicles and associated systems and methods
US8678321B2 (en) 2009-06-15 2014-03-25 Blue Origin, Llc Sea landing of space launch vehicles and associated systems and methods
WO2014177589A1 (en) * 2013-04-30 2014-11-06 Johannes Reiter Aircraft for vertical take-off and landing with hinged and bendable wings
US9896196B1 (en) * 2014-05-21 2018-02-20 Kaiser Enterprises, Llc Manned and unmanned aircraft

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