Methods and apparatus for effecting recovery of a high speed aircraft from a condition of incipient or developed spin

A method of effecting recovery of a high speed aircraft from a condition of incipient or developed spin involves, upon entry into the condition, deploying at least one parachute coupled only to the forebody of the aircraft forward of the center of gravity, and is particularly effective in recovering from a flat spin. After recovery, the chute or chutes can be collapsed, released or retracted. The chute is deployed in response to predetermined angle of attack and yaw rate of the aircraft, which may be accomplished automatically under the control of spin condition sensing means. The chute is highly effective because of its location, and hence may be relatively small, the size of the chute, when deployed being such as to have a negligible influence on the free falling rate of sink of the aircraft. The chute or chutes themselves may be conventional, typically of the pilot chute size employed to extract the main chute from a personnel parachute pack, and can be mounted in packed condition on the forebody fuselage or within the fuselage.

FIELD OF THE INVENTION 
This invention relates to spin control and recovery apparatus for use in 
high speed aircraft for effecting recovery of the aircraft from a 
condition of incipient or developed spin, and to methods of effecting such 
recovery, so as to regain control of the aircraft when it is in a spin. In 
particular, the invention relates to such apparatus and methods which are 
particularly suited for use on prototype full-scale spin demonstration 
fighter aircraft, and involves at least one parachute coupled only to the 
forebody of the aircraft forward of the center of gravity. The invention 
is, however, useful in production aircraft for spin recovery. 
BACKGROUND AND SUMMARY OF INVENTION 
All fighter aircraft projects are contractually required to perform a 
full-scale spin demonstration program. The purpose of the flight test 
program is to identify the types of spins that could be encountered 
inadvertently during future operational use of the aircraft, and the 
control techniques that are required to return the aircraft to the normal 
flight regime. Other purposes for these programs could include the 
demonstration of aerodynamic configurations that are spin resistant, 
automatic spin avoidance and/or recovery techniques, etc. In any event, 
spin demonstration aircraft are required now and will be required in the 
future, and these aircraft must be equipped with an emergency recovery 
system that is guaranteed to terminate any otherwise unrecoverable spin 
mode that might be encountered. 
One spin mode to which modern fighter-type aircraft are susceptible is the 
flat spin, wherein the aircraft exhibit "spinning top" motions. This type 
of spin usually has a high rate of rotation, an angle of attack between 
70.degree. and 90.degree., and effectively no spin radius, the aircraft 
spinning about an axis that passes through or near the center-of-gravity 
of the aircraft. Because an aircraft usually cannot recover from a 
developed flat spin through manipulation of the available aerodynamic 
controls, it is the spin which pilots fear most. 
To maintain a flat spin or any other type of spin, the aircraft must 
balance the nose-down aerodynamic pitching moment with an opposing 
gyroscopic pitching moment. The magnitude of the aerodynamic pitching 
moment is a function of the aircraft configuration, dynamic pressure and 
angle of attack (usually increasing progressively up to 90.degree. angle 
of attack). The magnitude of the gyroscopic pitching moment is a function 
of the mass distribution and the product of the roll and yaw rates 
generated about the aircraft body axis. In a flat spin, the yaw rate is 
considerably greater than the roll rate. The yaw rate required for spin 
equilibrium is determined by the magnitude of the aerodynamic pitching 
moment and the aircraft mass distribution. The other requirement for spin 
equilibrium is that the aerodynamic yawing moment about the body yaw axis 
be zero (actually very slightly propelling, i.e., pro-spin) at this yaw 
rate. Obtaining a flat spin requires, therefore, that a propelling 
aerodynamic yawing moment be generated at yaw rates below that required 
for balancing the aerodynamic pitching moment and that the magnitude of 
this yawing moment decrease (approaching a zero value) as the required yaw 
rate is attained. If a damping (anti-spin) yawing moment is generated 
below and at the required yaw rate the flat spin cannot be maintained. 
Emergency recovery systems used to date to generate an anti-spin yawing 
moment are complex, and usually incorporate a tail chute which is 
extremely inefficient when installed on modern aircraft that spin flat. In 
some instances, the chute size which is required for a particular type 
aircraft becomes impracticably large. In addition, the length of the riser 
line that attaches the tail chute to the aircraft is critical. If the 
riser line length is too short, the chute tends to collapse in the low 
dynamic pressure and reversed flow field that exists above the aircraft. 
If the riser line length is too long, the chute trails the aircraft at an 
angle which results in a nose-down pitching moment but no anti-spin yawing 
moment. Even the optmum riser line length results in a chute trail angle 
that contributes only a small antispin yawing moment. To compensate for 
the small anti-spin yawing moment, large parachutes are used. However, the 
use of large chutes results in off-design loads on the aircraft, which 
necessitates extensive internal and external reinforcement of the 
fuselage. In some cases, the reinforcement of the fuselage incurs changes 
in the mass distribution and external shape of the spin demonstration 
aircraft which jeopardizes the applicability of the results obtained from 
the testing. 
National Aeronautics and Space Administration has published a survey of 
spin-recovery devices and techniques, including rockets and wing-tipped 
mounted parachutes, but with particular emphasis on approaches in the 
design of tail-mounted spin-recovery parachute systems, including a 
compilation of design considerations applicable to spin-recovery parachute 
systems. The survey is published as NASA Technical Note D-6866, entitled 
"Summary of Design Considerations for Airplane Spin-Recovery Parachute 
Systems", by Sanger M. Burk, Jr., published in August, 1972. The 
disclosure of that publication is herein incorporated by reference, since 
many forms of apparatus disclosed therein are usable in my invention, even 
though my invention involves a basically contrary approach. The 
publication is also noteworthy in that it evidences the complexity of the 
problems involved and the great deal of effort which has gone toward their 
solution. 
Reference is here made also to my copending application, Ser. No. 570,505, 
filed Apr. 22, 1975, for Aerodynamic Spin Control Device for Aircrafts, 
wherein I disclose and claim a different solution to the same general 
problem, involving the use of doors or like devices mounted in the 
forebody of the aircraft so as to be deployable outwardly under spin 
conditions so as to change the fuselage yawing moment from pro-spin to 
anti-spin by changing the flow field over the fuselage forebody at high 
angles of attack, by effectively changing the geometric characteristics of 
the forebody. 
As indicated in the NASA publication referred to above, emergency spin 
recovery parachute systems have been traditionally attached to the aft end 
of the fuselage, although wing-tip-mounted parachutes have been given 
consideration, with notable lack of success. The tail-mounted technique 
resulted in a reasonably efficient recovery system for aircraft developed 
during the 1930-l950 time period. Since then, however, radically different 
airframe-propulsion configurations have evolved for which I consider this 
system to be unsuited. The efficiency of this technique progressively 
deteriorated over the years as the total plan form area increased, the 
distance between the center-of-gravity location and aft end decreased, and 
a greater proportion of the mass became concentrated in the fuselage. As 
the efficiency eroded, the complexity, size, and structural attachment 
load of the recovery installation escalated correspondingly. For some 
recent designs, the required structural modification costs have reached an 
absurd level. 
The major components of a conventional emergency spin recovery system 
installed at the aft end of the fuselage are: a deployment motor or gun, 
pilot chute, bridle line, deployment bag, riser line, ring slot or 
ribbon-type main chute, chute inflating device, and a chute tiedown and 
release mechanism. This emergency system is usually deployed during a flat 
spin, which is the most difficult spin to terminate, this being the case 
since the spin rate is the highest achievable while the aerodynamic 
controls required for recovery are operating at their lowest 
effectiveness. Also, the use of the controls may be lost completely on 
some aircraft after an engine flame-out is experienced. Unfortunately, 
when the need is greatest for the emergency system it is grossly 
inefficient since by virtue of its location on modern aircraft, which spin 
about a vertical axis passing through the center of gravity, only a small 
portion of the chute load produces the anti-spin yawing moment required 
for overcoming the spin rotation, whereas most of the chute load applies a 
nose-down pitching moment which generates a gyroscopic pro-spin yawing 
moment on current aircraft which have more of their mass concentrated in 
the fuselage than along the wing. Obviously, attachment of the recovery 
chute to the aft end of a high performance aircraft does not employ the 
laws of nature in an optimum manner but attempts to oppose them instead. 
The inefficiency of the tail-mounted chute location dictates the need for a 
large chute, which incurs a need for a major structural beef-up 
(accomplished internally and/or externally), an undesirable change in 
external lines, an undesirable increase in the inertia and change in the 
mass distribution, a bulky installation requiring special packing, 
inspection, etc. to keep size to a minimum, and a need for separating the 
chute from the aircraft before regaining normal flight. It should be noted 
that in one instance a 48 feet diameter chute attached to a 130 feet riser 
line was found to be inadequate. Also, the chute location is inefficient 
on other grounds which dictate additional severe design requirements. For 
example, the chute and riser line must be protected from heat before, 
during and after deployment of the system. The system must be designed to 
ensure minimum contact with aircraft structure, that is, avoid fouling of 
pilot chute and cuts and abrasion on riser line. The chute must trail in a 
stable position (not oscillate) to ensure an anti-spin yawing moment. This 
dictates the need for a high porosity type chute which requires an 
inflation device. The riser line length is critical, therefore, hopefully 
chosen properly since the main chute will collapse in the low q (dynamic 
pressure) reversed flow field above the aircraft if the line is too short. 
If it is too long, no anti-spin yawing moment is generated. 
The consequences of the described inefficiencies are high cost, long 
lay-ups of the test vehicle, system complexity, and a need for extensive 
(yet unsatisfactory) checkout system tests for evaluating a configuration 
which no longer represents the production vehicle. In addition, chute 
redundancy is not feasible, nor can installation on a production aircraft 
be considered. 
My invention incorporates a very small pilot type chute or chutes which can 
be stored on or in, and attached to, the nose of the aircraft, and which 
can be deployed therefrom. This device is highly effective because the 
chute location in operation takes advantage of the aircraft 
characteristics inherent in a high performance aircraft. For example, at 
deployment the total chute load applies a side force through a large 
moment arm and, consequently, a significant aerodynamic anti-spin yawing 
moment. As the rotation rate decreases and the chute realigns itself with 
the relative wind, a portion of the chute load applies a normal force and 
therefore a nose-up pitching moment which creates a gyroscopic anti-spin 
yawing moment. The chute operates in a relatively free-stream environment. 
Since the nose chute is extremely efficient, the chute size and 
corresponding load are very small, Therefore, no structural beef-up or 
external modifications are required, and no weight, inertia or mass 
distribution change need be incurred. This location also allows the use of 
an unstable chute. A low porosity chute can therefore be employed which in 
turn minimizes the required size and volume and needs no inflation device. 
The efficiency can be further enhanced through the use of sequenced chutes 
which are preprogrammed in an optimal fashion. A sequenced chute system, 
therefore, employs chutes which are smaller in size than would be required 
for a single chute system. 
In keeping with my invention, the parachute is deployed upon entry into a 
condition of incipient or developed spin, which deployment may be by the 
pilot upon entry into the spin, or in the incipient spin phase in response 
to a predetermined angle of attack and yaw rate sensed by a conventional 
air data computer, or automatically under the control of a spin condition 
sensing means carried by the aircraft. Although the chute size is 
variable, as one typical measure of size it may be said that the size of 
the chute, when deployed, is such as to have a negligible influence on the 
free falling rate of sink of the aircraft. For most aircraft, a maximum 
chute diameter of less than 10 feet is appropriate. When mounted or 
coupled forward of the cockpit, the dimensions of the chute should be such 
that the deployed chute can be collapsed against the side of the fuselage 
during a spin recovery dive without interfering with pilot visibility. The 
chute itself, or the chutes, may be considered as conventional, and may be 
mounted externally on the fuselage in a conventional chute bag, or may be 
housed in a special housing inside the fuselage and deployed therefrom in 
what may be considered a conventional manner. 
I am aware that there have been previously proposed various arrangements of 
parachutes in the nose portion of an aircraft, although for completely 
different purposes, these purposes being typically to hopefully lower an 
aircraft gently to the ground after a power failure. A typical example is 
found in Frost U.S. Pat. No. 2,673,051, issued Mar. 23, 1954, wherein a 
large supporting parachute is housed in the nose portion of a propeller 
driven VTOL aircraft so as to be deployable during vertical takeoff or 
landing in the event of a power failure. As a supporting parachute, the 
parachute would have to be necessarily quite large, and there is no 
relationship between the concept of this arrangement and my invention. A 
further typical example is found in Krahel U.S. Pat. No. 2,352,721, issued 
July 4, 1944, wherein large parachutes are coupled to the nose and other 
portions of the aircraft so as theoretically to lower the aircraft to the 
ground and prevent a crash. British Pat. No. 1,057,362, published Feb. 1, 
1967 (McMahon and Brown) discloses a parachute arrangement for lowering a 
rocket to the ground, the parachute being coupled to the forebody of the 
rocket and to the tail portion. These examples are conceptually remote 
from my invention, but they constitute the only instances known to me 
wherein a parachute is coupled to the forebody of an aircraft, regardless 
of the purpose. 
Other features and advantages of my invention will be set forth in or 
apparent from the following detailed description of presently preferred 
embodiments, taken in conjunction with the attended drawings.

DESCRIPTION OF PREFERRED EMBODIMENTS 
Referring now to the drawings, particularly FIGS. 1 and 2, the forebody of 
a high speed aircraft is indicated generally at 1, the illustrated 
aircraft being a spin demonstration or test aircraft. Its center of 
gravity is shown at 6, with the canopy shown at 2. Reference character 3 
designates a radome housing located forward of the first structural 
bulkhead 4. The instrument nose boom is shown at 5. Because of the very 
small size of the required chute or chutes, one or more sequenced solid 
flat type pilot chutes can be packaged and mounted externally in front of 
the cockpit windshield on the test aircraft as indicated at 8, with the 
secured suspension lines 9 extending forwardly and having their connecting 
"D" ring 10 attached to a known closed jaw-type clamp, which in turn is 
secured to the aircraft where the instrument boom enters the fuselage. In 
this manner, the chutes can be secured to the aircraft externally on the 
ground, and the status of the chutes and attachment verified at all times 
by the pilot. Additionally or alternatively, chutes can be mounted 
internally in a cylindrical container 11 and the "D" ring 12 attached to a 
closed jaw-type clamp secured to the first structural bulkhead 4 located 
in the nose. Thus, in test aircraft either or both arrangements may be 
used, but for production aircraft application, the internally housed 
arrangement would be used. 
The spin recovery device is activated, for example, by the pilot by opening 
the chute bag and/or compartment after closing the latch of the jaw-type 
clamp and thereby locking the chute attachment mechanism. This can be 
accomplished in a number of ways which are not particularly unique to the 
device described herein, as will be discussed in more detail in connection 
with the insuing figures. FIG. 1 shows one chute comprising canopy 14 and 
suspension lines 13 in its initial open position after deployment, the 
spin direction being indicated by arrow 7. FIG. 2 illustrates the final 
position of that chute, with the relative wind vector being shown by arrow 
15. 
FIG. 3 shows an illustrative attachment and release mechanism whereby the 
suspension lines are secured through their "D" ring to the aircraft. The 
suspension lines are shown at 13, and their "D" ring is coupled to a 
shackle 22 by a pin 23. Shackle 22 is in turn secured to the aircraft by a 
movable jaw member 19 which is pivotally mounted at 20, and held in the 
illustrated position by shear pin 21. A latch member 17 is pivotally 
mounted at 18 for latching and unlatching the movable jaw 19 through a 
control member 16 which extends to the cockpit. Latch 17 is shown in solid 
lines in its open position and in broken lines in its closed latched 
position. Latch 17 is in its open position and jaw 19 is in its closed 
position as shown for take-off, landing, and normal flight operation to 
allow for premature jettisoning of the parachute should it open 
accidentally. Thus, if premature parachute deployment occurs, the 
parachute load will shear pin 21, and permit jaw 19 to pivot clockwise so 
as to release pin 24, and hence shackle 22 and the parachute. For normal 
deployment of the parachute, latch 17 is moved to its closed position upon 
or prior to spin entry so as to retain the parachute in the jaws, thus 
coupled to the aircraft. The illustrated latched mechanism is not novel, 
and is fully disclosed on Page 52 of the previously mentioned NASA 
publication. 
FIG. 4 illustrates diagrammatically a typical external housed chute 
package. This may comprise a typical flapped personnel chute bag 8 
attached to the skin of the fuselage 1, and being closed by overlapping 
inner and outer flaps 8a and 8b, these flaps being secured in the closed 
position by a perforated cone-like member 8c coupled to the inner flap 8b 
and passing through an opening in the outer flap 8a, with a metal latch 
pin 8d passing through the perforation in the cone-like member to secure 
the bag 8 in closed position. Pin 8d is attached to the end 26 of a 
shielded release cable 27 which extends to the cockpit. The bag 8 includes 
a first chute 8e which is deployable by a spring capsule 25, and 
subsequently deployable second and third chutes 8f and 8g. When the flaps 
are unsecured by removing pin or pins 8d through release cable 26, the 
flaps open, and chute 8e is deployed by spring capsule 25. In a known 
manner, chute 8e automatically deploys chute 8f at a point in time which 
produces the most optimum recovery effectiveness, this being accomplished 
with a static line of a predetermined length that is attached to the 
canopies of the first and second chutes. This static line is shorter than 
would be required to allow the first chute to be fully inflated, and the 
specific length is selected to control the integrated aerodynamic effect 
both chutes contribute to spin recovery. The third chute 8g, as well as 
any additional chutes, are sequentially deployed in the same manner by 
attaching further static lines in the same manner as between the first and 
second chutes. FIG. 5 diagrammatically illustrates a typical internally 
housed chute arrangement, the chutes being housed in a cylindrical 
container 11 within the fuselage of the aircraft, and normally closed by a 
hinged cover 28, secured by a latch shown generally at 29, and which is 
controlled by a release cable 30 which leads to the cockpit. 
The chutes usually will collapse against the side of the fuselage during 
the spin recovery dive. Because there are and need be no riser lines, and 
since the suspension lines are short (because of the small diameters of 
the chute canopies), the collapsed chutes do not block the engine inlets 
nor do they interfere with pilot visibility. The collapsed chutes, 
however, are visible to the pilot. In any event, my invention permits the 
pilot the following three options after spin recovery: leaving the 
collapsed chutes outside the fuselage, releasing the chutes by opening the 
jaw clamp, or reeling in the chutes. 
The latter option, that of reeling in the chutes, is achievable through 
incorporation of a rotating drum or reel 31 as diagrammatically indicated 
in FIG. 6. In this figure, the releasable jaw clamp is schematically 
indicated at 19, coupled to the "D" ring 12 of a chute. The chute includes 
canopy 14 and suspension lines 13, and additionally includes a tether line 
33 which is coupled to the center of the canopy, and passes therefrom 
downwardly through a guide and pulley 34 secured to ring 12, with the 
lower end 32 of tether line 33 being secured to drivable drum or reel 31. 
This reel or drum may be located anywhere inside the fuselage nose, and 
the tether line appropriately guided thereto. The tether line is of 
sufficient length to permit the chute to be completely inflated during 
spin recovery. As the tether is partially reeled in, the collapsed chute 
load is further reduced as the canopy of the chute is pulled inside-out 
and completely deflated. If desired, the chute can be reeled in completely 
by appropriately arranging the drum 31 and the internally packaged chute 
configuration. 
FIG. 6 also indicates a static line 35 extending from the illustrated chute 
for coupling to a sequentially deployable chute. 
It will be understood that normally the externally and internally housed 
chute will not be used together, although there is no reason why they 
cannot be so used. Typically, though not necessarily, the external package 
arrangement will be used for spin demonstration or test aircraft, whereas 
the internally housed chute package is appropriate for production 
aircraft. Particularly for the external package, a deploying spring 
capsule is not necessary, since centrifugal and aerodynamic forces are 
entirely adequate to deploy the chute or chutes from the bag. 
The operation of the latch mechanism of FIG. 3 can be accomplished manually 
by the pilot or automatically, and the same is true for the deployment of 
the chute or chutes. 
The basic concept of my invention involves the location of the spin 
recovery chute and the efficiencies and advantages which accrue therefrom, 
and it will be appreciated that this concept can be carried out with 
numerous forms of apparatus different from that specifically disclosed 
herein. Therefore, although the invention has been described with respect 
to exemplary embodiments thereof, it will be understood by those or 
ordinary skill in the art that variations and modifications may be 
effected within the scope and spirit of the invention without departing 
therefrom.