Ejection sequencing system with airspeed and altitude sensing

An ejection sequencing system used particularly in an ejection seat operation includes continuous sensing of both airspeed and altitude. The continuous airspeed sensing is accomplished by continuously measuring the airstream temperature and then determining the optimum instant for parachute deployment from the measured temperature, thus making the deployment event temperature-responsive. When the sensed airspeed is within a predetermined acceptable limits, a first signal is generated. Continuous altitude sensing is accomplished by an aneroid barometer and when the sensed altitude is within predetermined acceptable limits, a second signal is generated. When these first and second signals are transmitted to the parachute container, they initiate parachute deployment. Redundant, failsafe system operation can be provided through the use of two or more interconnected airspeed sensing devices having fixed time delays for backup.

This invention relates to a system for deploying a parachute during 
ejection of an occupant from a disabled aircraft. More particularly, this 
invention relates to a system by which deployment of the parachute (which 
will lower the ejected occupant safely to the earth) is at least partially 
controlled by means of a true airspeed sensing mechanism responsive to the 
static and total temperatures of the airstream into which the occupant is 
ejected. 
Ejection seats are often provided in aircraft, particularly military 
aircraft, for assuring that the occupant will be able to escape from the 
aircraft in an emergency situation, whether such situation is occasioned 
by aircraft failure or by combat conditions. In either event, once the 
seat is ejected from the aircraft, the manseat combination continues along 
its trajectory in the airstream for a predetermined time after which the 
recovery parachute is deployed out of its pack for lowering the man safely 
to the ground. It is obvious that an ejection can occur under varying 
types of conditions, such as high altitude or low altitude and high speed 
or low speed. Within these widely varying conditions, it is extremely 
desirable that some means be provided for automatically deploying the 
parachute at the earliest appropriate time. 
In an ejection seat system, there are two basic parameters which govern the 
appropriate time for parachute deployment in any ejection situation. 
First, the true airspeed existing at the time the parachute is deployed 
must be beneath a certain limit, that limit being the one which produces 
opening forces which are within the structural capabilities of the 
parachute itself and which are physiologically tolerable for the ejectee. 
Second, deployment must not occur until the ejectee is below a certain 
altitude, that altitude being one at which a human being is capable of 
surviving for any protracted period of time. Ordinarily, this is 
considered to be an altitude of 15,000 feet. 
Additionally, in an ejection seat system employing a means for sensing 
airspeed and altitude after ejection, some fixed time delay must be 
incorporated into the system to allow the man-seat combination to leave 
the airflow surrounding the aircraft and to enter the true free airstream. 
This time delay provides sufficient time for the airspeed and altitude 
sensing system to monitor and to react to the environment of the free 
airstream. As a result, this time delay is in the range of 0.1 to 0.2 
seconds for a typical ejection operation, which would place the man-seat 
combination some 5 to 10 feet away from the aircraft and in the free 
airstream. 
In such an ejection seat system with on-board airspeed/altitude sensing, it 
is necessary to actually measure the airspeed during the time interval in 
which the time delay is taking place. In the past, airspeed sensing 
systems have been of two general types, both of which employ pitot tubes, 
but such sensing systems have not proved to be altogether satisfactory. 
The common pitot tube system is an aerodynamic pressure measurement system 
which measures "equivalent" airspeed, working on the assumption that the 
altitude or density is that of sea level. However, parachute performance 
under conditions of widely varying air density for the most part is 
determined by "true" airspeed as opposed to "equivalent" airspeed and 
since the range in which the parachute may have to operate varies from sea 
level to 15,000 feet, this type of sensing system introduces an error in 
true airspeed which can be as great as 21% which is practically 
unacceptable. The other way commonly used for measuring airspeed is to use 
a true airspeed indicator and machmeter which measures dynamic pressure as 
well as static pressure and ambient temperature. While a system of this 
type is quite satisfactory for use in connection with airplanes which are 
ordinarily flying in a straightforward path at a particular altitude, 
their suitability for use in connection with an ejection seat is poor, 
because such systems are very complex and unreliable in the ejection seat 
environment. Pitot tubes are normally sensitive to their alignment with 
the airstream to such a degree that if the angle of the pitot tube with 
respect to the airstream is greater than 30.degree., the accuracy of the 
measurement is severely degraded. Thus, the use of a special pitot probe 
is required to extend the maximum value of this allowable probe angle to 
airstream angle to 60.degree. or greater, without degrading system 
accuracy to an unacceptable value. 
Also in the past all airspeed sensing and altitude sensing ejection seat 
systems operated on the basis of selecting fixed time delays, depending on 
the airspeed being above or below a predetermined value and the altitude 
being below 15,000 feet. These fixed time delays were selected either at 
the instant of ejection seat initiation or within 0.3 second after 
seat/aircraft separation. With these fixed time delays it was impossible 
to optimize the system sequencing for all ejection airspeeds. Each time 
delay was determined by the maximum ejection airspeed and altitude 
combination and it was hence too long for ejections occurring at any speed 
or altitude other than the maximum airspeed, maximum altitude combination. 
With the foregoing in mind, it is, therefore, an object of the present 
invention to overcome the difficulties and deficiencies associated with 
the prior art and to provide instead a new and improved ejection seat 
sequencing system, incorporating continuous sensing of both airspeed and 
altitude. 
Another object of the present invention is to provide means for combining 
the outputs of multiple sensors so as to achieve redundant, failsafe 
system sequencing to assure highly reliable successful system operation 
with either optimum or near optimum system sequencing even in the event of 
an airspeed sensor failure. 
Another object of the present invention is to provide a novel airspeed 
sensing system wherein the airspeed is measured on the basis of the 
airstream static and total temperatures, which, of course, are not 
dependent upon the particular orientation of the man-seat combination in 
the airstream. 
Other objects, advantages and salient features of the present invention 
will become apparent from the following detailed description, which, taken 
in conjunction with the annexed drawings, discloses a preferred embodiment 
thereof. 
The foregoing objects are attained by providing an ejection seat sequencing 
system which includes a continuous airspeed sensing system which measures 
only two variables, namely, the free-stream total and static temperatures. 
Once these two temperatures are known, the true airspeed can be calculated 
by means of an equation and the equation itself can be solved by an 
electrical or electronic circuit. Such a circuit continuously measures the 
difference in the voltages arising from the difference in the sensed total 
and static temperatures, and compares that value with a reference voltage. 
When the value of the differential voltage reaches the value of a 
reference voltage, an electrical signal is transmitted to an altitude 
switch operated by a pressure device such as an aneroid bellows. In this 
manner, the electrical circuit determines when the airspeed is proper and 
the aneroid bellows determines when the altitude is proper, and when both 
proper conditions have been met, the appropriate signal is transmitted 
from the altitude switch to commence deployment of the parachute. 
In the apparatus of the present invention, there are provided two or more 
airspeed sensors physically located at different points on the ejection 
seat, each having an electrical or electronic circuit operating as 
previously described. In addition, these circuits are interconnected in 
such a way as to require that both airspeed sensing systems indicate that 
the seat/occupant has reached the desired airspeed before either of the 
two circuits transmit an electrical signal to either or both of two 
altitude switches. 
The output of the airspeed sensors may be electrical, mechanical, or other 
form suitable for functioning in the existing environment. Since different 
techniques exist for measuring static temperature of the airstream as well 
as total temperature, all such means are not set forth herein, and 
instead, only representative devices which comply with the operating 
principles of this present invention are discussed.

Referring now to the drawings in further detail, there is illustrated 
herein in FIG. 1, airspeed sensing apparatus in accordance with the 
principles of the present invention, such apparatus being generally 
designated 10. The apparatus includes an airstream probe in the form of an 
elongated tubular housing member 12 having an open forward end 14 which is 
smoothed and rounded in configuration at its edge as shown at 16. An 
internal cavity 18 is provided at the forward end of the housing such 
cavity 18 gradually diminishing in cross-sectional diameter from the inlet 
end 14 to an elongated restricted throat portion 20. A static temperature 
sensing device 22 is positioned in the throat 20 and the electrical leads 
24 of this sensing device extend rearwardly through the remainder of the 
housing 12. 
The rear end of the housing 12 is designated 26. A cylindrical bore 28 is 
provided in the rear portion of the housing extending from the rear end 26 
toward the restricted throat portion 20. A tapering section 30 is provided 
where the bore 28 merges into the throat 20. 
An interior housing member 32 is disposed centrally within the bore 28 and 
is retained therein by means of supporting vanes 34 connected between its 
forward end and the tapering portion 30 of the bore and rear vanes 36 
connected between the rear end of the interior housing 32 and the rear end 
of the main housing 12. The interior housing 32 is provided with a central 
internal chamber 38 within which is disposed another temperature sensing 
device designated 40. Electrical leads 42 from the temperature sensing 
device 40 extend through the rear end of the interior housing 32. 
The interior housing 32 is provided with an inlet orifice 44 at its forward 
end and an outlet or exit orifice 46 at its rear end. The diameter of the 
inlet orifice, as shown, is at least twice as great as the diameter of the 
exit orifice. The forward end of the interior housing 32 is rounded or 
tapered as shown at 48 to facilitate airflow around it. The inlet and 
outlet orifices 44 and 46 are aligned centrally within the apparatus 10 
and hence are in alignment with the central axis of the throat 20. 
The construction of the probe 12 of FIG. 1 is such that airflow from the 
airstream will enter through the inlet 14, pass through the internal 
cavity 18 and into the throat or restricted portion 20 where the 
temperature of such air is sensed by the sensor 22. The air then passes 
from the throat 20 into the bore 28. Such air passes through the vanes 34 
and 36 and exhausts through the outlet end 26 of the housing. However, a 
portion of the air passing through the bore 28 enters the auxiliary 
housing 32 through the inlet 44 therein. This portion then passes into the 
internal chamber 38 where the temperature thereof is sensed by the sensor 
40 and then exhausts through the outlet 46. The chamber 38 in the 
auxiliary housing forms a sheltered area where few if any normal airflow 
vectors are present. 
Temperature sensing by means of the apparatus of FIG. 1 is used to 
calculate airspeed in accordance with the following formula: 
##EQU1## 
wherein: V = true airspeed (feet per second) 
T.sub.t = free-stream temperature (degrees Rankine) 
T = ambient temperature (degrees Rankine) 
.gamma. = ratio of specific heats 
g = gravity 
R = universal gas constant 
Since it is known that for air .gamma. = 1.4, g = 32.2 ft/sec and R = 53.3, 
the formula then becomes: 
EQU V = 109.6 .sqroot.T.sub.t - T [2] 
the temperature sensing device 22 serves as the free-stream temperature 
sensor for measuring T.sub.t while the temperature sensing device 40 
serves as the ambient temperature sensor for measuring T. The electrical 
leads 24 and 42 connect these sensors with an electric circuit as is shown 
in FIG. 2. 
Temperature sensing devices 22 and 24, which are illustrated as temperature 
responsive resistors, are connected in series circuit relationship with 
calibrating resistors 50 and 51, respectively, in two legs of a bridge 
circuit, the other two legs of which are formed by fixed resistors 52 and 
53 of equal value. The junction between resistors 52 and 53 is connected 
to a Zener diode-regulated source of d.c. voltage +V and the junction 
between resistors 22 and 40 is connected to a point of lower potential at 
the junction between a Zener diode 54 and a fixed resistor 55, this Zener 
diode and resistor forming a regulated divider between the positive d.c. 
source and ground. 
The remaining corners 56 and 57 of the bridge circuit are connected through 
fixed input resistors 58 and 59, respectively, to the input terminals of a 
conventional differential amplifier 60, the output of which is connected 
through an output resistor 61 to the gate electrode of a silicon 
controlled rectifier (SCR) indicated generally at 62. A feedback resistor 
63 is connected between the output and one input terminal and a feedback 
resistor 64 is connected between the other input terminal and a point of 
reference potential which is formed by a Zener diode reference circuit 
including Zener diodes 65 and 66 and a resistor 67 connected in series 
circuit relationship with each other and in parallel circuit relationship 
with a battery 68. A second point of reference potential between diode 66 
and resistor 67 is also connected to the amplifier to establish the 
operating level thereof. The end terminals of a potentiometer 70 are 
connected between the +V source of voltage and the cathode of SCR 62, the 
movable wiper of potentiometer 70 being connected to the gate of the SCR 
to establish an initial operating level adjustment. The output voltage is 
developed across a load resistor 71 which is connected between the cathode 
of SCR 62 and the junction between diode 54 and resistor 55. 
The operation of the circuit shown in FIG. 2 is as follows. When the 
resistance exhibited by sensing resistors 22 and 40 is equal, the bridge 
is balanced and no voltage appears between bridge corners 56 and 57, 
resulting in no output signal. However, when a sufficient difference 
between the temperature sensors 22 and 40 exists in a preselected 
direction, a difference in potential between points 56 and 57 is 
developed, producing a voltage difference at the input terminals of 
differential amplifier 60, which, being a high gain amplifier, produces an 
output signal through resistor 61 to the gate of SCR 62. The amplifier 
operating level and its output polarity is such that the SCR is held in a 
non-conductive state by the output signal of the amplifier. As the 
airspeed of the ejected seat decays and the temperature difference of the 
temperature sensors 22 and 40 also decays, when the desired airspeed is 
reached the amplifier output becomes such to render the SCR conductive, 
producing an output signal across load resistors 71. As previously 
indicated, this output signal is transmitted to an altitude switch. 
Referring now to FIG. 3, there is designated therein in block diagram form 
an altitude switch generally designated 76 which is connected with the 
sensor circuit of FIG. 2, such circuit being generally designated 78. The 
output signal from the circuit 78 is transmitted to the altitude switch 
76. An aneroid barometer generally designated 80, of a conventional 
design, is also connected with the altitude switch 76. The purpose of the 
aneroid barometer is to sense when the altitude is below a predetermined 
magnitude and altitude pressure. The inlet to the aneroid barometer is 
sheltered or shielded so that no dynamic pressure forces are sensed. 
Hence, the sensor circuit 78 senses the airspeed and provides an output 
signal to the altitude switch 76 only when the airspeed is within the 
predetermined safe range. The aneroid barometer 80 senses the altitude and 
provides an output signal to the altitude switch 76 only when the altitude 
is below its preset value. Once both signals have been properly 
transmitted to and received by the altitude switch 76, such switch closes 
to transmit a signal to a parachute container generally designated 82 for 
commencing deployment of the parachute therein. 
Referring now to FIG. 4, there is shown therein an ejection seat system 
with redundancy and failsafe features. The aircraft ejection seat 
generally designated 100 carries two complete systems and two probes, 
which are labelled Airspeed Probe -1 and -2. For convenience of 
illustration, the systems and seat are drawn in duplicate but it should be 
understood that both systems are in actuality employed on a single seat. 
The Aircraft Power passes through an Automatic Connect/Disconnect to a 
Connector on the seat and from there to a normally open Mechanically 
Actuated Switch. The Seat Firing Controls are mechanically connected 
through a Safe/Arm Handle and a Ground Safety Pin connection to the 
Mechanically Operated Switch. In preparation for flight the Ground Safety 
Pin is removed and the Safe/Arm Handle is moved to the Armed position. 
Thus, when the Seat Firing Controls are actuated in flight, the ejection 
sequence is initiated by mechanically closing the Mechanically Operated 
Switch and energizing the Mechanically Operated Thermal Battery which 
requires some 50 to 60 milliseconds to get up to power. 
Meanwhile, once the Mechanically Actuated Switch is closed, the aircraft 
power passes through a connector to initiate the Catapult Ignitor and the 
powered inertia reel device, called P.I.R.D. Igniter. It then passes 
through the Automatic Connect/Disconnect to the Canopy 
Jettison/Fragilization which either breaks or jettisons the aircraft 
canopy. The operation of the inertia reel holds the ejection seat occupant 
in position while the catapult starts ejecting the seat/occupant 
combination or ejected load from the aircraft. At a point in this ejection 
operation, the aircraft power is disconnected from the seat. 
After the thermal battery has come up to rated output, it supplies the 
electrical power as required by the 0.36, 3.0, and 1.4 second delay 
elements and by the Airspeed Sensor Switch. Each airspeed probe connects 
to a sensor switch and the two Aircraft Sensor Switches are themselves 
interconnected, as shown by the cross arrow between them, so that if 
either switch is open, both are open. That is, if either sensor senses a 
high speed condition, both systems are blocked from further operation. 
As soon as the Airspeed Sensor Switch receives information from the probe 
indicating that the true airspeed is below a predetermined maximum amount, 
such switch closes and sends a charge to the Capacitor. If the ejected 
load is then beneath 14,000 feet, the 14,000 Foot Altitude Switch is 
closed and the signal from the capacitor is transmitted directly to the 
Parachute Container Opener to open the container to commence parachute 
deployment while at the same time releasing the drogue parachute by the 
signal sent to the Drogue Release/Transfer. If the ejected load is above 
the 14,000 foot level, the 14,000 Foot Altitude Switch will remain closed 
thus preventing parachute deployment. By having the Airspeed Sensor Switch 
charge the Capacitor, sufficient power is available at whatever time the 
load reaches 14,000 feet, even if the thermal battery has already stopped 
functioning. 
As can be noted, there are three time delays in the circuit of FIG. 4. The 
0.36 Second Time Delay is needed for the sensor to sense the airspeed and 
come to equilibrium. The other two time delays are safety delays to assure 
that the parachute will be deployed satisfactorily even if both airspeed 
sensors malfunction. Thus, after 1.4 seconds, power is sent through the 
1.4 Second Delay Switch to a 7,000 Foot Altitude Switch. If that switch is 
closed, meaning that the ejected load is at an altitude of less than 7,000 
feet, the parachute container is immediately opened, regardless of whether 
the Airspeed Sensor Switch has operated. Similarly, after 3 seconds, power 
is sent through the 3.0 Second Delay Switch and Capacitor to the 14,000 
Foot Altitude Switch. If that switch is closed, meaning that the ejected 
load is at an altitude of less than 14,000 feet, the parachute container 
is immediately opened, regardless of whether the Airspeed Sensor Switch 
has operated. 
It should be apparent that various obvious changes and modifications can be 
made to the system disclosed herein without departing from the spirit and 
scope of the invention as defined in the appended claims.