Polaris guidance system

An inertial guidance system for a ballistic missile that is completely independent of ground support at and after time of launch comprising:

• stable platform means for providing space-stabilized reference,
• triad of accelerometers for providing acceleration measurements of said stable platform as it moves through space, said accelerometers requiring no information signals from sources external of the missile nor any airborne computation of the position of the missile in space;
• triad of gyros for providing angular measurements of said stable platform as it moves through space, said gyros requiring no information signals from sources external of the missile nor any airborne computation of the position of the missile in space,
• said triad of gyros positioned on said stable platform means so that the sensitive axis of each gyro is aligned coincident with one of coordinate axes X, Y, and Z of said stable platform, said X and Y axes being substantially in a horizontal plane at the time of launch and said X and Z axes being in a vertical plane that is substantially in the plane of the desired trajectory,
• said triad of accelerometers positioned on said stable platform means so that the sensitive axis of each is aligned coincident with one axis of an orthogonal X', Y', and Z' reference frame of said stable platform, said Y' axis being coincident with said Y axis, and said X' and Z' axes being in a vertical plane that is substantially in the plane of the desired trajectory, and said X' and Z' axes being respectively above and below said horizontal plane by approximately forty-five degrees,
• first, second and third driving means electromechanically intercoupling said gyros and said stable platform means for moving said stable platform about inner, middle, and outer axes of a stable platform gimbal system,
• first, second, third, fourth, and fifth resolver means mechanically coupled to said first, second, and third driving means,
• said first driving means coupled for movement about the inner gimbal axis in response to signals from a gyro positioned with its sensitive axis along said Z axis, and said second driving means and said first resolver means coupled together for movement about the inner gimbal axis in response to signals from gyros positioned with their respective sensitive axes along said X and Y axes, and said third driving means and said first resolver means coupled together for movement from the outer gimbal axis in response to signals from gyros positioned with their respective sensitive axes along said X and Y axes,
• output means from said second resolver means coupled to electrically transmit signals representative of angular movement about said inner gimbal axis to the inputs of said third and fourth resolver means,
• output means from said third resolver means coupled to electrically transmit a signal depending on angular movements about said inner and middle gimbal axes to the input of said fourth resolver means,
• first amplifier means electrically coupled to receive an output signal from said fourth resolver means that is representative of movement of the missile about its roll axis, the sensitivity of the roll angle signal being independent of the missile's pitch angle, and further being directly representative of roll error,
• second amplifier means electrically coupled to receive an output signal from said fourth resolver means that is representative of movement of the missile about its yaw axis, the sensitivity of the yaw angle signal being independent of the missile's pitch angle,
• third amplifier means electrically coupled to receive an output signal from said fifth resolver means that is representative of movement of the missile about its pitch axis, said pitch axis being coincident with the outer gimbal axis,
• computer means coupled to receive the signal outputs of said triad of accelerometers that are representative of acceleration of said stable platform, respectively, along said X', Y', and Z' axes, said computer means processing the acceleration signals against preset parameters to provide control signals along the pitch and yaw axes of the missile, and an engine cut-off signal,
• pitch coupling means receiving signals from said computer means and said fifth resolver means for combining these respective signals into a composite signal representative of pitch error, and
• yaw coupling means receiving signals from said computer means and said fourth resolver means for combining these respective signals into a composite signal representative of yaw error.

An inertial guidance system for a ballistic missile that provides navigation and control while it is proceeding in a flight trajectory to a predetermined target, comprising:

• a first inertial reference frame comprising mutually orthogonal axes X, Y and Z, said Y axis in a horizontal plane at the time of launch, said X and Z axes in a plane that is perpendicular to said horizontal plane at the time of launch and substantially in the plane of said trajectory, and said X axis being elevated 45.degree. above said horizontal plane and substantially in the direction experiencing the highest dynamic range of input acceleration,
• a second inertial reference frame comprising mutually orthogonal axes X', Y' and Z', and said X' axis parallel to said X axis, and said Y' axis parallel to said Y axis, and said Z' axis parallel to said Z axis,
• an inertial measuring unit comprising a triad of accelerometers, the sensitive axis of each accelerometer aligned with one axis of said first reference frame, and said accelerometers providing signals representative of linear motion of said vehicle, said accelerometers requiring no information signals from sources external of the missile nor any airborne computation of the position of the missile in space, and said measuring unit further providing a triad of gyros, the sensitive axis of each gyro aligned with one of said axes in said second reference frame, said gyros providing signals representative of angular motion about the respective axis of said second reference frame, said gyros requiring no information signals from sources external of the missile nor any airborne computation of the position of the missile in space,
• a digital computer coupled to receive said signals representative of linear motion for processing against preset parameters stored in said computer for providing control signals along the pitch and yaw axes of said ballistic missile, and
• autopilot means coupled to receive said control signals from said computer means and said signals representative of angular motion from said gyros for providing mechanical control movements.

An inertial guidance system for a ballistic missile comprising

• a triad X, Y, and Z of inertially fixed mutually orthogonal reference axes, and at the time of launch of said missile, said X axis elevated by substantially 45.degree. above the horizontal, said Z axis in the direction approximately 45.degree. below said horizontal, and said X and Z axes substantially in the plane of the intended trajectory,
• first, second and third accelerometers, the sensitive axis of each accelerometer coinciding with said X, Y, and Z axes, respectively,
• first and second function generators, and a tan .alpha. multiplier, said multiplier receiving the output of said first accelerometer,
• first, second and third summing means, said first summing means adding the negative of the output of said first function generator to the negative of the output of said first accelerometer, said second summing means adding the negative of the output of said second function generator to the negative of the output of said second accelerometer, and said third summing means adding the negative of the output of said tan .alpha. multiplier to the output of said third accelerometer,
• first, second and third time integrators, said first integrator integrating the added output signal of said first summing means and incorporating in the integration the initial velocity-to-be-gained magnitude in the X direction, said first integrator thereby generating the velocity-to-be-gained signal along said X direction which is the input to the first and second function generators, said second integrator integrating the added output of said second summing means, and said third integrator integrating the added output of said third summing means, and incorporating in the integration the initial velocity-to-be-gained in said Z direction, and a negative Z command neutralizing signal magnitude,
• first, second, third and fourth amplifiers, said first amplifier amplifying the output signal of said second integrator, said second amplifier amplifying the output signal of said second accelerometer, said third amplifier amplifying the output of said third integrator, and said fourth amplifier amplifying the output of said third summing means, and
• fourth and fifth summing means, said fourth summing means subtracting the output signal emitted by said first amplifier from the output of said second amplifier, the difference constituting the yaw rate command signal for said missile, said fifth summing means adding the output signals of said third and fourth amplifiers and the sum constituting the pitch rate command signal of said missile.