There’s a lot of circumstantial evidence. How many Rolling Airframe Missiles was General Dynamics working on in 1975, when the patent was filed, would GD be patenting things and doing exploratory work before landing the contract? (Surprisingly two actually, the FIM-92 and what would become the RIM-116, which happen to both share a Seeker).
The biggest smoking gun linking them together would be, the presence of the Feedback loop in the block diagram(s), being introduced as a Difference (damping) mechanism via Pulse Width Modulating the power (and polarity) provided to the Torque motor.
( US4037806 ) The servo error signal, which is the difference between the control servo command and the wing position feedback signals, is superimposed on a dither oscillator signal, causing the servo switching circuit to pulse modulate power to a torque motor. The torque motor integrates this incoming pulse modulated signal so that the resultant wing deflection rate follows the servo error signal.
Happens to have a Stinger (note the four forward surfaces not two, as with the the Redeye), and further describes Shoulder launch, which the RIM-116 regardless of intent at 162lbs The RAM would require a very determined Soldier, let alone the requisite launcher.
Relevant diagrams
( US4054254 ) The pitch rate sensor illustrated in FIG. 6e is also responsible for the damping or decreasing of the output of the control surface incidence. The high initial pitch rate reflected by the trace 86 results from the high initial deflection from the control surfaces. In time, the output of the pitch rate sensor reduces the undamped control signal and produces a stabilized pitch rate as is suggested by the portion 88 of trace 86.
Since the movements of the control surfaces have been timed to correspond to the coincidence of the control plane with the direction toward the flight path intended (vertically upward), there is an acceleration of the vehicle in the earth related upward direction as is illustrated in FIG. 6c by the trace 90. It will be noted that the trace reaches a maximum level in approximately one and one-half revolutions and sustains that level throughout the duration of the control excursions with very little over-shoot. The corresponding trace 92 in FIG. 6d for the earth related horizontal plane shows that substantially all of the acceleration is in the direction of intended change for the flight path. When the input command terminates, as is illustrated in FIG. 6a at point 94, a wing incidence sequence substantially the reverse of that occurring when the command signal commenced is initiated. This wing incidence sequence is illustrated in FIG. 6f as portion 84 of the trace 82. It will be noted that the maximum signal now corresponds to the inverted position of the control plane and therefore, causes the vehicle to pitch back towards its original flight attitude. Since the input command terminated the wing incidence is almost solely a function of the acceleration signal 98, this acceleration signal then is the equivalent of static stability such as is utilized in open-loop control in maintaining a constant flight attitude in the absence of a control signal. The portion 95 of the pitch rate signal 86 in FIG. 6e reflects response of the pitch rate sensors to the opposite angular velocity. The effect of pitch rate damping summer 54 is to enhance the acceleration signal at this portion of the control sequence, to thereby produce a damping of the commanded wing incidence and as a result, to prevent over-shoot. It will be noted that only a minimal overshoot of the vertical acceleration as evidenced by portion 96 of FIG. 6c is experienced. The wing incidence and pitch rate damp out in the next five revolutions and reach zero at approximately the same point, representing the return of the airframe to a zero-angle-of-attack, stable-flight, mode.
