Abstract: An example reusable high-altitude balloon system includes a balloon with a first end supporting a payload and a second end with an aperture and an apex fitting that is positioned within the aperture. A clamp applies a pressure to a plurality of pleated folds formed in the perimeter of the aperture around the apex fitting to form an air-tight seal against the balloon at the perimeter of the aperture. The reusable high-altitude balloon system further includes control circuitry that controllably releases the apex fitting from the balloon to initiate a descent sequence.

Abstract: Provided is a radar and communications enhanced sail for a sailboat, sail ship, or sail drone. The sail includes a first sail section comprising an active communication system, a second sail section comprising a passive communication system, or a combination thereof. The active communication system includes an antenna array (transceiver) and a software-defined radio (SDR), while the passive communication system comprises a reflective panel or sections and/or array of reflector panels or sections. The active system utilizes its SDR and transceiver to communicate back and forth with an onshore SDR and transceiver to provide information as necessary. The passive system receives a radar signal via the reflective material on the sail and reflects the signal back at the radar, which produces a radar cross section indicating that there is an object (in this case the sailboat) in the ocean.

Abstract: The technology relates to techniques for drogue deployment for a lighter than air (LTA) vehicle descent. A drogue deployment system for an LTA vehicle descent can include a drogue comprising a drogue parachute coupled to a carrier. A spring can be configured to launch the drogue from a launch tube directed outward from an apex of the LTA vehicle in an acute angle from a horizontal plane. A core can be placed around the launch tube and placed around the spring, the core compressing the spring and holding the spring in a compressed state prior to deployment, and a riser can couple the carrier to the envelope of the LTA vehicle. In some cases, the drogue deployment system can comprise two or more drogues, wherein intervals between the two or more drogues can be selected such that horizontal components of drogue deployment forces approximately cancel out.

Abstract: An example reusable high-altitude balloon system includes a balloon with a first end supporting a payload and a second end with an aperture and an apex fitting that is positioned within the aperture. A clamp applies a pressure to a plurality of pleated folds formed in the perimeter of the aperture around the apex fitting to form an air-tight seal against the balloon at the perimeter of the aperture. The reusable high-altitude balloon system further includes control circuitry that controllably releases the apex fitting from the balloon to initiate a descent sequence.

Abstract: The technology relates to techniques for drogue deployment for a lighter than air (LTA) vehicle descent. A drogue deployment system for an LTA vehicle descent can include a drogue comprising a drogue parachute coupled to a carrier. A spring can be configured to launch the drogue from a launch tube directed outward from an apex of the LTA vehicle in an acute angle from a horizontal plane. A core can be placed around the launch tube and placed around the spring, the core compressing the spring and holding the spring in a compressed state prior to deployment, and a riser can couple the carrier to the envelope of the LTA vehicle. In some cases, the drogue deployment system can comprise two or more drogues, wherein intervals between the two or more drogues can be selected such that horizontal components of drogue deployment forces approximately cancel out.

Abstract: An example reusable high-altitude balloon system includes a control system configured to initiate a termination sequence by separating a payload from an Earth-facing end of the balloon body. Separation of the payload from the balloon body generates a torque that causes the balloon body to invert and release lift gas through a vent duct, initiating a flight termination sequence that facilitates landing of the reusable balloon system without destroying the balloon body.

Abstract: A laminated load ring for a balloon assembly includes a plurality of ring stacking units stacked one on top of the other. Each of the ring stacking units can include a main body having a central opening, a plurality of arms, and at least one weld line. The plurality of arms may each extend away from the main body around a circumference of the main body. The at least one weld line can be formed on the main body. The plurality of arms of the plurality of ring stacking units may be aligned with one another. The weld line of each of the plurality of ring stacking units may be offset from the weld line of a directly adjacent ring stacking unit in a direction extending around the circumference of the laminated load ring.

Abstract: An aerospace vehicle comprising an airplane or spacecraft, operatively coupled to an airship balloon containing lighter than air gas adapted to elevate the vehicle. A control system adapted to deflate the balloon upon reaching a specific altitude by directing the gas to a propulsion system and high pressure gas chambers for powering the vehicle at a greater speed or to a greater altitude. The balloon can be retracted into the vehicle to achieve a better aerodynamic shape and further re-inflated for decreasing the speed of the vehicle upon reaching a destination.

Abstract: A lift assembly for use during inflation of a balloon envelope is provided. The lift assembly includes a plate structure that has a set of cavities. Each cavity includes one or more openings passing through the plate. One or more pistons are coupled to the plate through at least one of the openings of each cavity in the set of cavities. Each piston has a hollow tube portion projecting lengthwise from the at least one opening, a flange attached to the hollow tube portion and a grabber portion in communication with the flange. The grabber portion includes a plurality of bearings for grabbing a stud attached to an apex of the balloon envelope. A handle portion is coupled to the plate. The handle is arranged to lift the balloon envelope when the bearings have grabbed a given stud.

Abstract: The present disclosure provides a method and apparatus for turning an envelope of a balloon into a parachute. The balloon may include a payload, an envelope filled with a lift gas, and a parachute system. The parachute system may include a fuse couple to the envelope. The parachute system may also include an activation system coupled to the fuse. The activation system may be configured to ignite the fuse. The fuse may be ignitable to melt through at least a portion of the envelope to separate an upper portion of the envelope from a lower portion of the envelope. The upper portion of the envelope may be coupled to the payload such that when separated from the lower portion, the upper portion performs as a parachute for the payload.

Abstract: An apparatus is provided that includes a solenoid chamber having a plunger configured to create an opening in a first balloon envelope of a balloon system when the solenoid chamber is actuated. The opening enables gas to evacuate from the first balloon envelope. A mixer valve having a nozzle is coupled to the solenoid chamber. The mixer valve is configured to create a gas mixture formed from the gas evacuating from the opening and atmospheric gas. A heating device is attached to the nozzle. The heating device is configured to heat the gas mixture created by the mixer valve. A second balloon envelope is attached to the nozzle. The second balloon envelope is arranged to expand in proportion to a quantity of the heated gas mixture passing through the nozzle. This expansion of the second balloon envelope creates an amount of lift for controlling descent of the balloon system.

Abstract: A balloon includes a cut-down device, a payload, and an envelope. A control system could be configured to determine a position of the balloon with respect to a predetermined zone. The cut-down device could be operable to cause at least the payload to land in response to determining that the position of the balloon is within the predetermined zone. The predetermined zone includes an exclusion zone and a shadow zone. The shadow zone could include locations from which the balloon would be likely to drift into the exclusion zone based on, e.g., historic weather patterns or expected environmental conditions. Boundaries of the shadow zone could be determined based on, for example, a probability of the balloon entering the exclusion zone.

Abstract: A method and aircraft system are provided for determining an optimal gliding speed that maximizes gliding distance of the aircraft upon engine failure, and comprises calculating the optimal gliding speed from a speed polar curve, true airspeed, heading, and vertical wind speed. The polar curve may be constructed with consideration of the aircraft center of gravity and weight. The calculated gliding speed, which may be restricted to a threshold, is provided to the pilot and optionally to an autopilot. The method is repeated periodically, or if the wind vertical or horizontal velocities exceed a threshold for a predetermined period of time.

Abstract: A balloon launch unit produces lift gas to enable balloon launch or to allow the prolongation of balloon flight. A method of using the balloon launch unit can include reforming a fuel source by reaction with water to generate hydrogen-rich lift gas mixtures, and injecting the lift gas into a balloon. A reforming operation can include causing the combustion of a combustible material with ambient oxygen for the release of energy; and heating a reforming combination of reaction fuel and water with the energy released from the combustion of the combustible material, to a temperature above that required for the reforming reaction wherein the fuel and water sources are reformed into lift gas. The amount of the combustible material combusted is sufficient to result in the release of enough energy to heat an amount of the reforming reaction fuel and water sources to the temperature above that required for the reforming reaction.

Abstract: A navigation system for spinning projectiles using a magnetic spin sensor to measure the projectile roll angle by sensing changes in magnetic flux as the projectile rotates through the earth's magnetic field is disclosed. The magnetic spin sensor measurements are used to despin a body reference frame such that position, velocity, and attitude of the projectile can be determined by using a strapdown inertial navigation system (INS) algorithm. More particularly, a multisensor concept is used to measure pitch and yaw angular rates, by measuring Coriolis acceleration along the roll axis and demodulating the pitch and yaw rates therefrom.

Abstract: The recoverable airborne instrument platform accurately determines its present position and uses this data to execute a predetermined flight plan and ultimately guide its descent to a predetermined landing site. This is accomplished by installing the instrument package payload in the aerodynamic exterior housing of the recoverable airborne instrument platform, which has a plurality of moveable control surfaces thereon to autonomously control the altitude, attitude and flight path of the recoverable airborne instrument platform. A navigation circuit contained within the aerodynamic housing determines the geographic location of the recoverable airborne instrument platform as well as the location of at least one predetermined recovery site. The determined position data is used to dynamically calculate a flight path which allows the guidance control circuit to both execute a predetermined flight plan and controllably descend the recoverable instrument platform to a selected predetermined recovery site.

Abstract: A missile guidance system employs a strap down inertial measurement unit with a separate accelerometer and star sensor that are moved for preflight accelerometer initialization. During flight, light is directed on the accelerometer and inertial measurement unit to produce on the star sensor images of the light source, the accelerometer and the inertial measurement unit. The location of these images, which manifest the acceleration and position of the inertial measurement unit relative to the vehicle, are used to improve the accuracy of the inertial measurement unit data for vehicle guidance.

Abstract: A system for protecting a pendulous oscillating gyroscopic accelerometer (POGA) from power failure interruption (PFI) events, the POGA having a servo driven member (SDM), rotor driven member (RDM), and torque summing member (TSM), wherein the POGA is a part of an accelerometer system of a guidance system which detects such events and communicates them to the accelerometer system. In the invention, the acceleration is stored just prior to the PFI induced system shut-down, and after the PFI the POGA drives are turned back on. The velocity change during the PFI is determined based on the TSM deflection occurring during the PFI event, acceleration existing at the beginning, and elapsed time. This is reported to the guidance system in a timely fashion.

Abstract: For an exact vectoring of active or combat projectiles, a GPS-supported pilot projectile is provided with a satellite navigation system or GPS receiver and continuously detects its own position and speed during the complete operational (firing) sequence--and thus also the ballistic interference variables--and transmits this information via a transmitter to the ordnance unit for evaluation and correction of the firing parameter of a subsequently fired active or combat projectile. If the active or combat ammunition or projectile to be vectored is equipped with submunitions, then the GPS receiver and the transmitter are located in the pilot submunition, so that the ballistic interference variables affecting the submunition can be recorded as well.

Abstract: The present invention relates to a floatable aircraft such as a parachute balloon. The parachute balloon is comprised an envelope which is able to be collapsed and stored during periods of nonuse. A parachute is attached to the envelope, enabling the parachute balloon to fall safely by the effect of the parachute in various circumstances including the usual descent or an emergency during a fall, such as when out of fuel.

Abstract: A method for determining the initial conditions for an inertial measurement nit (IMU) of a second vehicle launched from a wing of a first vehicle is provided. The method includes the steps of defining a state vector x as including (a) the rotation .zeta. of the computed coordinate axes with respect to the real coordinate axes of the second vehicle and (b) the projection .delta..alpha. along the Z axis of the first vehicle of the rotation of the second vehicle from its nominal coordinate axes to its real coordinate axes. A measurement z is defined as the projection .delta..beta. of a rotation angle .beta., along the Z axis of the first vehicle, between the nominal coordinate axes and a current computed coordinate axes. The method also includes the steps of estimating x over time with a Kalman filter, wherein the projection .delta..beta. is the measurement vector and the state vector x changes only due to random noise and processing x to produce the attitude about the Z axis of the first vehicle.

Abstract: An aircraft of the present invention has an envelope used in such an air balloon and/or airship. The aircraft can be in flated continuously to the envelope with supplying an air into the envelope without extending the envelope on the ground, and the envelope can be rolled up continuously into the gondola with an inside-out condition, as keeping the inflating condition of the envelope.

Abstract: An attitude sensing safing system for a missile, which senses an attitude ror of the missile, and activates a dudding signal when the error exceeds a certain value for a predetermined time. The missile includes a delay arming timer, which is deactivated when by the dudding signal. The system allows the missile to correct itself, by requiring that the error signals persist for a set time.

Abstract: A guidance monitoring safety system for a guided missile having platform os and electromechanical integrators with inputs from a guidance information analyzer. The system monitors malfunctions resulting from erratic or unsafe missile flight either dudding the missile if any malfunctions are present or providing a safe-to-arm fuzing signal in the absence of malfunctions.

Abstract: A multi-node Lambert Guidance System that controls missile velocity along a missile flight path through adjustment of missile attitude during guidance of the missile through an intercept point and one or more way points. The system utilizes a linear Lambert solution in that it is derived from a plurality of Lambert solutions each corresponding to a particular node along the prelaunch flight path solution and each including corresponding time varying weights on the accuracy at several flight path nodes. The present invention allows direct translation of guidance velocity corrections into vehicle attitude adjustment commands, and thereby avoids additional computation required in prior predictive integration guidance approaches.

Abstract: As an improvement to the architecture, utility, efficiency and cost associated with the procurement, setup, and operation of a real-time tracking/autonomous destruct and data recovery system for vehicles capable of exhibiting great velocities, accelerations, and attitude rates during all mission phases. A vehicle based independent tracking system 10 consists of an on-board vehicle data tracking system 12, an on-board vehicle data transmission system 14, an off-board flight data monitoring system 16, and a remote computer setup/test system 18. An interface umbilical-a 82 interconnects on-board vehicle data tracking system 12 with on-board vehicle data transmission system 14 for real-time transmission of dynamic time, space and position information of rockets and like vehicles to off-board flight data monitoring system 16 during all mission phases via an RF communication interface 86.

Abstract: A missile is tested by first stimulating the performance of missile built-in tests that are stimulated in service by signals from the launch site, and which tests do not require internal access to the missile. If performance is satisfactory, no further testing is required. If a condition of unsatisfactory missile performance is detected, an access cover is removed, connections are made to internal sources of data, and the built-in tests are repeated in order to determine which internal components caused the unsatisfactory missile performance. The component is repaired. The testing with the access cover removed is repeated, and, if the performance is satisfactory, the access cover is replaced and the testing with the access cover present is repeated.

Abstract: A dual flywheel assembly for use in an airborne vehicle for storing mechanical energy therein prior to flight and for permitting the extraction of electrical energy therefrom during flight, which assembly includes two flywheels which are linked by a suitable linkage structure such that, if roll motion of the vehicle starts to occur during flight, the flywheels tilt in equal and opposite directions out of their normal planes of rotation, which tilting motions act in a passive manner to stabilize the roll motion of the vehicle.

Abstract: Missile motion information from an inertial measurement set and the angle-to-target information from a seeker/sensor is used to estimate range and/or time to go on a homing missile without a direct measurement of range or range rate. Up to six comparisons of spectrally matched linear and angular measurements, corrected by target state estimates and analytical coupling terms, are used to determine the error in the range estimate. A nonlinear sorting and filtering procedure rejects the unusable error sources among the six potential error sources, and thresholds against too little information. The stability margin in the missile guidance loop is used as a sensor indicator for excessive estimated range.

Abstract: A system for controlling the placement of a gun launched projectile adapted with a radial impulse motor incorporated into the projectile for imparting a thrust subsequent to launching and a receiver incorporated into the projectile for targeting information for determining the projectile trajectory and the projectile roll rate, roll position, and pitch, i.e., vertical reference; and a computer linked to the receiver and the radial impulse motor for determining the projectile trajectory and the projectile roll rate, roll position, and vertical reference time after launch of the projectile and angle of corrective vector to ignite the radial impulse motor to affect the trajectory of the projectile to land the projectile on a desired target.

Abstract: A guidance system for internally controlling the flight path of a missile includes a guidance platform having dispersion control means mounted on the guidance platform for detecting acceleration of the missile due to lift and side forces. The dispersion control means includes dispersion detection means to calculate the velocity and position errors relative to a drag-only trajectory from the detected acceleration due to the lift and side forces. The internal guidance system also includes missile positioning means for controlling the position of the missile platform to substantially eliminate the velocity and position errors. The guidance system also includes level finding means for determining a substantially horizontal level axis and the orientation of the missile relative to a level axis. The missile positioning means also performs other functions such as missile leveling in which the wings-level axis of the missile is aligned to the level axis.

Abstract: Process for the autonomous positional control of guided missiles, which are to be launched from a carrier plane, but without the cooperation thereof, wherein data on the instantaneous position of the guided missile, as well as flight and target data are fed into a computer of the guided missile already on the ground. The initial data fed into the computer are continually updated with changing values from the inertial navigation and additional navigational aids and are further calculated to achieve autonomous flight guidance and self-guiding of the missile. An uninterruptible power supply of the guided missile is maintained beginning from the moment of data reception.

Abstract: One or more strap-down star tracker modules and a strap-down gyro module cooperate to provide a high-performance and low-cost inertial guidance system especially well suited for ballistic missiles. The one or more star tracker modules include an alignment laser that cooperates with one or more gyro module autocollimation mirrors to measure three-axis bending and twist between the one or more strap-down star tracker modules and the strap-down gyro module. The one or more strap-down star tracker modules and the strap-down gyro module respectively determine inertial angle independently of one another. The strap-down gyro module is compensated for gyro drift whenever any difference remains between the inertial angle measured thereby and the same inertial angle measured by the one or more star tracker modules after isolating the contribution thereto of the measured three-axis bending and twist.

Abstract: A friction compensation system which augments a stabilization rate loop includes a first summer (11) responsive to a rate command signal and a feedback signal to provide a torque command signal, a second summer (17) responsive to the torque command signal via torquer (15) and the gimbal torque disturbances to provide a net applied torque, a gimbal (19) subjected to the gimbal torque disturbances and having an inertial rotation governed by the net applied torque and a gyro (21) responsive to the inertial rotation of the gimbal to provide the feedback signal. An adaptive friction compensator responsive to the torque command signal provides an enhanced torque command signal to which an intermediate third summer (33) is responsive in accordance with a predetermined model (31) which is adjusted adaptively in realtime. The third summer is responsive to the torque command signal and the enhanced torque command signal to control the gimbal.

Abstract: A system and method that uses differential computation of position relative to a global positioning system (GPS) coordinate system and the computation of an optimum weapon flight path to guide a weapon to a non-moving fixed or relocatable target. The system comprises an airborne platform that uses a navigation subsystem that utilizes the GPS satellite system to provide the coordinate system and a synthetic array radar (SAR) to locate desirable targets. Targeting is done prior to weapon launch, the weapon therefore requires only a navigation subsystem that also utilizes the GPS satellite system to provide the same coordinate system that the platform used, a warhead and a propulsion system (for powered weapons only). This results in a very inexpensive weapon with a launch and leave (autonomous) capability. The computational procedure used in the platform uses several radar measurements spaced many degrees apart. The accuracy is increased if more measurements are made.

Abstract: The system may be used, for example, to initialize the inertial unit of a missile from the inertial unit of a vehicle carrying the missile. According to the invention, a servo-control circuit is provided for slaving data in the missile inertial unit on data from the carrier inertial unit, together with a Kalman filter connected in parallel with said servo-control circuit and designed to identify harmonization errors between the two inertial units.

Abstract: A method of determining the difference between the actual rotational speed of the body of a guided missile and a nominal rotational speed is shown to consist of the steps of comparing a shaped output of a Hall effect cell with a shaped output of an oscillator and deriving a signal indicative of any difference between the periods of the shaped outputs.

Abstract: A reaction-wheel stabilized spacecraft reduces attitude errors at wheel reversals by application of a dither component to the wheel torque command signal.

Abstract: An inertial transformation matrix generator generates a succession of Euler transformation matrices in inertial coordinates, and is useful for converting to inertial coordinates the responses of a sensor hard mounted on the hull of a craft (e.g., an aircraft). First, second and third rate-sensing gyros located proximately to said sensor are strapped down to the craft hull, and are oriented to sense the motion of the craft hull in three mutually orthogonal directions, for providing respective output signals indicative of components of craft hull motion in each of those three mutually orthogonal directions. The output signals of the gyros are digitized, and based on these digital signals successive incremental Euler transformation matrices are generated.

Abstract: An attitude and heading reference detecting apparatus having three gyros and three acceleometers mounted to a navigation vehicle in such a manner that their input axes are made coincident with principal three axial directions of the navigation object, respectively, a magnetic azimuth sensor, a signal transforming section supplied with the outputs from the three gyros, the outputs from the three accelerometers and the output from the magnetic azimuth sensor, a computing section supplied with the output from the signal transforming section and a signal output section supplied with the output from the computing section, for producing an attitude and heading reference signal and for delivering the same to the outside.

Abstract: An Earth-based, rapid-fire, electromagnetic accelerator system is provided for launching multiple hypervelocity nuclear or non-nuclear independently targetable warheads on ballistic trajectories to targets located anywhere on or above the Earth's surface. The warheads are mounted inside a reinforced launching sabot containing a plurality of coaxial superconducting dipole magnets. The sabot is magnetically accelerated to hypervelocities inside a large-bore vacuum tube by sequentially exciting a series of driving coils mounted coaxially along the tube. The sabots are injected into the tube from pre-evacuated storage canisters thereby eliminating the need for an air-lock. Terminal guidance systems allow the warheads to be fired over intercontinental distances to hit small, preselected targets with nearly perfect accuracy. The launch velocities are sufficiently high to enable warheads to also intercept and destroy orbiting satellites moving in space high above the Earth's surface.

Abstract: A method and apparatus for identifying inertial reference unit (IRU) errors in a guided missile employing a multi-mode guidance system and constructing correction terms to recover the missile true position. Discrepancy parameters are introduced to indicate misalignment between missile and launching platform (or ship) inertial frames where the missile onboard executive computer simultaneously processes the data provided from missile onboard sensors and target relevant data uplinked from the launching platform. The discrepancy parameters are employed to construct correction factors used to reduce the discrepancies. This updated missile configuration is then coupled with the target state estimator outputs to reconstruct smoothed line-of-sight (LOS) angles for terminal homing engagement.

Abstract: A strapped-down intertial system for a vehicle including computational means and a sensor package. The sensor package includes a first part (1) fixedly mounted on the vehicle, and a second part (5) which is rotatable about a first axis (Z). On the second part (5) are fixedly mounted gyroscopic means (13, 15) responsive to rotation about two axes, one of which is the first axis (Z), and accelerometer means (9, 11) responsive to acceleration along two axes, one of which is not the same as either of the two axes of the gyroscopic means (13, 15). The computational means is arranged to calculate the heading of the vehicle from the outputs of the gyroscopic means and accelerometer means taken with the first part (1) stationary, and the second part (5) successively in at least three different angular positions about the first axis (Z).

Abstract: A roll measuring device for missiles comprising an inertia member supported about an axis by a case and provided with means whereby rotation of the case relative to the inertia member about the said axis can be measured, the inertia member being balanced about the axis so as to be rotationally uninfluenced by gravity or movement medial or angular to the axes.

Abstract: An apparatus comprising, inter alia, gyroscope feedback circuitry which aw scanning the rotor magnet of a gyroscope disposed in an associated guided missile system so as to increase the seeker field thereof is disclosed. A signal from the cage coil of the gyroscope having an amplitude approximating a sine function of the angular position of the spin axis of the rotor magnet portion of the gyroscope with respect to the body axis of the associated guided missile is used to generate a constant amplitude drive signal for driving the precession coil of the gyroscope. Scanning, so as to drive the rotor magnet in a predetermined scan pattern is accomplished by phase shifting the signal from the cage coil as a function of its amplitude and then driving the precession coil with the aforementioned constant amplitude drive signal which is phased-locked to the phase shifted or processed cage coil signal.

Abstract: Mastless sails for a boat or ship which comprise sails having capability for utilizing static and dynamic lift in combination with drag, gravity, and line tension to provide stability and propulsion forces for the boat or ship, communication capability located near said sails when aloft for use for said boat or ship, sensing capability located near said sails when aloft for use by said boat or ship, and apparatus for raising and lowering the communication and sensing apparatus.

Abstract: An inertial guidance system for a rocket powered ballistic missile which vides navigation and control while the missile is proceeding in its flight trajectory. Signals are generated within the confines of the missile and without outside information representative of angular and linear motion of the missile with respect to an inertial frame of reference, a digital computer compares these signals against a preset parameter stored in the computer to provide control signals along the pitch and yaw axes. An autopilot uses these control signals to provide mechanical control movements, and an exhaust nozzle deflection arrangement connected to the autopilot deflects the exhaust of the rocket for guidance along a predetermined flight trajectory.

Abstract: An attitude control system is provided for a body-stabilized orbiting space vehicle having roll, pitch and yaw axes. Sensors carried by the vehicle provide signals representative of angular deviations about at least two of said orthogonal axes. An actuator for the system comprises two wheels rotatable about respective non-aligned spin axes fixed with respect to the vehicle. An electronic control circuit receives signals from the sensors and controls the rotation speeds of the wheels for attitude correction about two of said orthogonal axes. Short term stabilization about the third of the orthogonal axes is provided by the gyroscopic stiffness of the actuator.

Abstract: The invention comprises a novel strapdown gyro inertial system having a single degree of freedom between it and the vehicle upon which it is mounted. An angular transducer having a low friction bearing is mounted on the vehicle along the vehicle roll axis bearing and supports the strapdown gyros and accelerometers that are skewed with respect to the roll axis. The method of determining the total vehicle attitude employs the gyros so that the vehicle rate around each gyro axis is integrated and combined with the vehicle angular roll to obtain the total attitude. The inertial momentum of the sensor block, electrically enhanced or strictly dependent upon its mass, decouples high acceleration inputs along the free axis and thus reduces the dynamic rate sensing requirement along the free axis. Also the novel configuration of the sensors skewed with respect to the free axis, permits easy calibration of gyro scale factors and other inertial parameters which are characteristics of gyroscopes and accelerometers.

Abstract: An electronic guidance system which when provided with a set of target poion coordinates prior to launch, calculates the missile position and velocity in an inertially fixed coordinate system during flight. All inertial coordinate system variables are defined by a three axis orthorganal coordinate system centered on the target position. This guidance system functions without an active roll control by continuously pointing the velocity vector of the missile towards the target position.