DRONE AND SHIPBOARD GUIDANCE SYSTEM

The Navy Model QH-50D Drone (hereinafter referred to as the drone) is a remotely controlled, rotary- wing, weapon-carrying vehicle designed specifically for deployment from the deck of destroyer type ships in antisubmarine warfare (See figures 1 and 2).

The Target Control System AN/SRW-4B (hereinafter referred to as the shipboard guidance system) is used to control the drone in the accomplishment of its mission.

The drone and the shipboard guidance system, together, constitute the DASH (Drone Anti-Submarine Helicopter) Weapon System which is a basic feature of the Navy FRAM (Fleet Rehabilitation And Modernization) program.

The drone and certain components of the over-all system are manufactured by, or under the program management of, Gyrodyne Helicopter Company. The model T50- BO- 10 or T50-BO- 12 Turboshaft Engine (hereinafter referred to as the engine) and the remainder of the system components are items of government furnished equipment (GFE). Figure 3 of this section lists the principal components of the DASH Weapon System and figure 4 lists the leading particulars of the drone.

The drone is maintained stable in flight by the Automatic Flight Control Set AN/ASW-20 (AFC set). The Radio Receiving Set AN/ARW-78 (receiving set) receives and decodes the command signals which originate on the destroyer and superimposes them on the stabilization system. Primary electrical power for the airborne system is obtained from the airborne generator which is driven by gearing from the rotor drive system.

The digital command guidance, or data link, system is of the pcm/fm (pulse code modulation of a frequency modulated subcarrier) type, in which the commands are transmitted to the drone as discrete binary coded pulses.

The drone is capable of carrying either one of the two following weapon stores:

1. Two Mark 44 Mod 0 torpedoes. 
2. One Mark 46 Mod 0 torpedo.

The drone mission capabilities are described in "Operating Limitations” section under the paragraph headed OPERATING ENVELOPE. 

DASH WEAPON SYSTEM MISSION

    The drone is controlled from either one of two stations on the destroyer: one at the flight deck (figure 5), the other at the combat information center (CIC). The drone is housed in a hangar on the destroyer and is transported along the deck on the launcher-retriever restraint cables. Shipboard radar is used for tracking the drone, and sonar is used for the detection of underwater targets. A display system combines the radar and sonar information to enable the CIC controller to position the drone with respect to the target. A typical combat mission against an enemy submarine is described briefly in the following steps.

1. The fueled and armed drone is moved from the hangar and spotted at the launch area on the flight deck.

2. The drone is secured to the deck by means of a tiedown device incorporating a remotely actuated disconnect coupling, and the restraint cables are removed.

3. Using the control monitor, the deck controller starts the drone engine, and the rotors are brought to normal rpm by the engine limiter. The controller transfers power from the auxiliary power source to the drone generator, performs a brief prelaunch check, and actuates the umbilical cable remote disconnect.

4. The deck controller actuates the hold-down release, applies an altitude command, and maneuvers the drone away from the ship. He executes a smooth climb-out and vectors the drone in the direction of the target.

5. The deck controller transmits to the CIC controller, via interphone, the altitude, airspeed, and heading of the drone. The CIC controller sets this data into the CIC control. When the drone appears on the CIC radar display, an integrated transfer of control from deck to CIC Is effected.

6. The CIC controller plots the drone course, taking into account relative motion of destroyer and target, wind direction, and drone heading and speed. The CIC controller heads the drone on this course and monitors the closure on target, readjusting drone heading, ff necessary, to compensate for unpredicted changes.

7. As the drone approaches the target area, the CIC controller establishes the proper drone altitude, airspeed, and heading for weapon delivery. At the optimum torpedo drop point, the CIC controller releases the weapon or weapons. The CIC controller then commands the proper altitude and heading for the return to the vicinity of the ship.

8. When the drone comes into sight, control is transferred from the CIC controller to the deck controller who executes the approach and landing.

9. The deck controller stops the engine by remote control, the restraint cables are rigged, and the drone is returned to the hangar. 

DRONE

The drone (figures 1 and 2) is a helicopter of the coaxial rotor type. The two-bladed semi-rigid rotors rotate about a common center; the upper rotor turns counterclockwise, and the lower rotor turns clock- wise (viewed from above). Power is transferred from the turbo-shaft engine to the rotors through a two-stage transmission system. Engine speed is maintained within limits by a power turbine limiter and a fuel control unit.

The drone structure is of open construction and consists essentially of a transmission and fuselage support casting, tubular landing gear struts and skids, and the tubular aft fuselage assembly. The fuel tank and certain components of the AFC set are mounted on the aft fuselage. 

SHIPBOARD GUIDANCE SYSTEM

The shipboard guidance system provides means for controlling the drone from the destroyer. All components of this system (figure 3) are installed on shipboard.

The shipboard guidance system consists of certain basic components common to many target control systems, plus other components for specific use with the DASH Weapon System. These specific components are items 14 through 17 of figure 3. Items 18 through 27 of figure 3 may be used in other drone or target control systems. 

ENGINE

    The drone is powered by a Boeing Model T50-BO- 10 or - 12 two- shaft gas turbine engine, which consists of two major sections; a gas producer section, and a power output section. There is no mechanical connection between the rotor in the gas producer section and the rotor in the power output section. Their rotational speeds may very independently as required by the load.

The gas producer section consists of a two-stage axial and centrifugal compressor directly coupled to a single-stage axial flow turbine wheel, two combustion chambers, an accessory drive section, and a rotor housing and sump.

The power output section comprises an axial flow, single-stage turbine wheel, power output reduction gearing, accessory drive gearing, and an output shaft.

An interstage bleed valve is mounted on the underside of the bleed air collector, and is used to provide a means of discharging a portion of compressor air to atmosphere to prevent compressor surge during engine operation. The bleed valve is opened by a bellows and linkage assembly when the differential between interstage pressure and compressor pressure exceeds a preset value. The interstage bleed valve is installed on both T50-BO- 10 and - 12 engines; however, on some T50-BO-10 engines, due to variations in compressor internal configuration, the bleed valve is inactive. T50-BO-10 engines with activated bleed valves are easily identified by the presence of a pneumatic tube (compressor discharge pressure line) connecting the bleed valve to the compressor case. These engines are military rated at 330 shp at 90º F. T50-BO-10 engines with inactive bleed valves do not have the pneumatic tube installed, and have a nominal available horsepower rating of 300 shp at 90º F. All T50-BO-12 engines utilize activated bleed valves, and are military rated at 330 shp at 90ºF, and 365 shp at 59ºF.

The nameplate mounted on the accessories section housing of the T50- BO-10 or -12 engines shows the horsepower rating of the engine under standard ambient conditions (59ºF). The horsepower rating shown on the nameplate of T50-BO-10 engines with inactive interstage bleed valves is 330 shp. This rating is equivalent to a minimum rating of 300 shp at 90º F. T50- BO- 10 engines with activated interstage bleed valves have a horsepower rating of 365 shp shown on the nameplate. This rating is equivalent to a minimum rating of 330 shp at 90º F. Since all T50-BO-12 engines have activated interstage bleed valves, the horsepower rating shown on the nameplate of these engines is 365 shp (equivalent to a minimum rating of 330 shp at 90ºF).

ENGINE IGNITION SYSTEM

The engine ignition system consists of a relay, an ignition unit, two igniter plugs (one for each burner), and the necessary cabling.

The ignition relay provides means for energizing the ignition unit and igniter plugs. It is used only during the starting cycle.

The 28-vdc input, capacitor-discharge ignition unit provides a high voltage output for the igniter plugs during the starting cycle.

Each igniter plug has two electrodes, and is connected to the ignition unit. During the starting cycle, high voltage from the ignition unit is applied to the center electrode, which sparks to the ground electrode, igniting the fuel. After combustion occurs, the igniter plugs are de-energized; combustion is self-sustaining.

STARTER-GENERATOR

An air-cooled starter-generator is in constant engagement with the gas producer accessory drive gear train. During the starting cycle, the starter-generator is energized to drive the gas producer rotor until combustion has occurred. Generator output is not used in the present system.

ENGINE INSTRUMENTS

All engine-monitoring instruments are located in the control monitor. These instruments are disconnected from the engine when the umbilical cables are disconnected. The functions of the instruments are discussed under OPERATING CONTROLS AND INDICATORS. Instrument markings are discussed in Section V.

ENGINE FUEL SYSTEM

The engine fuel system includes a low-pressure filter, a fuel pump, a fuel control unit, a power turbine limiter, a fuel shutoff valve, a flow divider, two fuel nozzles, and interconnecting fuel and air lines.

The fuel control unit is modulated by the power turbine limiter, and in turn controls gas producer rotation to maintain the power output section at constant speed as required by engine loading. The power turbine limiter and the fuel control unit are inter- connected by a flexible air line.

The electrically operated fuel shutoff valve is used to shut the engine down safely from any operating condition. The valve is opened and closed electrically, but does not require any current to hold it in either position.

The flow divider is used to distribute fuel flow between the primary and secondary nozzle orifices in each burner assembly. Internal passages to the primary orifices are always open. When the fuel nozzle pressure reaches 150 psig, a valve in the flow divider opens, providing fuel flow through the secondary orifices. Engine fuel is MIL-T-5624, grade JP-5. The capacity of the fuel tank is 52 gallons (353.6 pounds).

ENGINE OIL SYSTEM

    The engine incorporates all lubrication system components necessary for engine operation under the climatic, altitude, and attitude conditions that fall within the engine operating envelope. The lubrication system consists of a supply sump, a lube and scavenge pump and pressure regulator valve, a filter bypass valve, an oil cooler and thermal and pressure relief valve, a magnetic chip detector, a torque sensing system, and interconnecting lines and passages. Lubricating oil in accordance with Specification MIL- L-23699 is used in the engine.

    Oil is drawn from the sump by the pressure element of the lube and scavenge pump, which pressurizes the system and supplies lubrication to the entire engine. The other two elements of the pump scavenge the rotor housing and output section. Drainage oil from other gas producer locations returns to the sump by gravity. System oil pressure is maintained by the internal relief valve. A full-flow type filter and filter bypass valve arrangement, which includes a disposable cartridge, is utilized in the engine. The bypass valve opens to assure adequate oil flow to the engine bearings when filter pressure drop exceeds 25 psi. A screen is located in the suction passage to each oil pump element. The oil cooler is mounted on the bottom side of the engine, and utilizes exhaust eductor air for cooling. Oil having temperatures below 71ºC (160ºF) bypasses the oil cooler by means of a thermostatically operated valve. The cooler also includes a bypass valve which opens when the pressure drop across the cooler exceeds 20 psi.

ROTOR SYSTEM

ROTOR ASSEMBLIES

    Rotary motion is imparted to each rotor by means of a saddle splined to each of the rotor shafts. The design and construction of the upper and lower rotor saddles and associated parts is similar but not identical. Axial motion of the saddles is prevented by spacers and a nut on the top of each rotor shaft. Pivotally mounted on each saddle are two plates which support the rotor spindles and tip brake bellcrank shafts. Pivotally mounted on the rotor spindles are the blade hub and pitch horn fittings which contain a series of radial and thrust bearings.  

    These fittings are secured axially on the spindles by positively locked retaining screws. Lugs are pro- vided on the outboard ends of the hubs to receive the rotor blades, which are provided with metal reinforcing plates at the root. A tip brake is pivotally mounted in the tip of each blade. The tip brake is connected to a crank arm on the tip brake shaft by a control rod, which runs through a tube bedded within the blade. The tip brake is opened by an inward pull on the control rod and is normally held closed during flight by centrifugal force acting on a counterweight on the brake. The control rod is connected to the tip brake through pin and slot connection to permit free outward motion of the control rod. The upper and lower rotor tip brake mechanisms are connected in such a manner that they move in relatively opposite directions. This permits either the upper or lower rotor tip brakes to be ex- tended at will without affecting the other.

Gust locks are provided for each rotor to prevent flapping or teetering while the rotors are static or rotating at low or medium speed during acceleration and deceleration. These locks are spring-urged into engagement and are disengaged by centrifugal force when approximately 75 percent of the full 610-rpm rotor speed is reached (approximately 460 rpm).

ROTATING CONTROLS

    Collective pitch and cyclic pitch control movements are applied to the rotor blades through swash plate assemblies on the rotor mast below each rotor (See figure 6 ). The design and construction of the two assemblies is similar but not identical, though they serve the same function. Each swash plate assembly consists primarily of an inner and outer ring with two single- row ball bearings between the rings. Each swash plate assembly is supported by a two-axis gimbal ring so that its plane of continuous rotation can be deflected longitudinally and laterally. The gimbal ring for the lower swash plate assembly is pivoted on a support, which is adapted for axial, but not rotative, movement on a cylindrical extension of the transmission bell housing. Axial motion is imparted by a collective pitch yoke which also restrains the inner swash plate ring against rotation (See figure 7). The inner swash plate ring is provided with two forked extensions: one aft of center, and one abeam to the left. Each fork is connected to an independent series of links and levers, which control the deflection of the plane of rotation of the outer swash plate ring through the non-rotating inner ring. Rotation is imparted to the outer swash plate ring, at the same rate as the rotor, by scissors linkages secured to an adapter, which is keyed to the rotor saddle.

    The upper swash plate assembly is gimbal-mounted on a support, which slides on, and rotates with, a sleeve, which is splined to the lower rotor shaft. Thus, the relative speed between inner and outer rings is twice the rotor speed. Four control links, 90 degrees apart, link the upper inner swash plate ring to the lower outer swash plate ring so that the same axial and tilting control movement is imparted to both swash plate assemblies simultaneously. A pitch link connects each rotor hub pitch horn to its associated outer swash plate ring. This pitch link leads the blade-feathering axis by 90 degrees.

    Collective pitch, which determines thrust (lift) for given rotor speed, is controlled by moving the swash plate assemblies axially on the rotor mast to cause all the rotor blades to rotate equally about their feathering axes. Tilting the swash plate assemblies causes a cyclic pitch change in each blade as it rotates. (See figures 8 and 9 ). Since the pitch links lead the blades by 90 degrees, the continuously varying pitch of each blade will cause the rotor disks to tilt in correspondence with the tilt of the swash plate assemblies. Thus, tilting the plane of rotation of the swash plate assemblies forward will cause the drone to assume forward speed. Lateral and rear- ward speed can be achieved by tilting the swash plate assemblies in the appropriate direction. Longitudinal and lateral tilt can be combined to cause flight in any direction with no change in heading of the fuselage. 

    The swash plates are automatically locked against cyclic pitch displacement during acceleration and deceleration of the rotors to prevent damage to the rotating control system. This cyclic control lock mechanism is mounted on the after side of the servo actuator and consists of two solenoid actuated pawls which engage notches in the longitudinal and lateral cyclic output arms of the servo actuator. The locks are engaged on the application of an engine-off command and are disengaged, after the rotors reach normal rpm, when power is transferred from the auxiliary motor-generator to the airborne generator. The control locks also may be actuated manually.

Fuselage heading is controlled by movable tip brakes at the tips of the rotor blades (See figure 10). At a given steady state heading of the fuselage, torque is applied equally to each rotor through the second stage pinion and upper and lower rotor shaft drive gears. 

Since the applied torques are equal and opposite, the differential torque is zero. When the tip brakes on one of the rotors are extended, the aerodynamic resistance of that rotor is increased, and a torque differential in the transmission is induced. This differential torque produces a change of fuselage heading. When the tip brakes are retracted, the zero torque differential is restored and the fuselage remains oriented at the new heading.

Since the upper rotor turns counterclockwise, and the lower rotor turns clockwise (viewed from above), extension of the upper rotor tip brakes causes the fuselage to alter heading to the right. Extension of the lower tip brakes causes a turn to the left.

    The linkage which controls the tip brakes includes an axial motion shaft running through the hollow center of the upper rotor shaft, a tip brake actuator horn, and links to the upper rotor tip brake bellcranks. Motion is transmitted to the lower tip brake bellcranks by means of struts from the horn to a swivel housing below the upper rotor. The swivel housing provides roller thrust bearing engagement with an axial motion sleeve splined to the lower rotor shaft and through the upper swash plate assembly. The lower tip brake actuator horn is secured to the lower end of the sleeve and is connected by links to the lower rotor tip brake bellcranks.

TRANSMISSION SYSTEM

    Power is applied from the engine to the rotors through the two-stage transmission. A quill shaft transfers power from the output section of the engine to the transmission first stage input pinion which meshes with the first stage gear (See figure 11). The first stage gear is mounted on, and splined to, the hub of the second stage pinion. The two second stage gears are mounted parallel to each other on a common center of rotation. Both mesh with the second stage pinion, one above and one below. Thus, rotation of the second stage pinion in a given direction causes the two-second stage gears to rotate in relatively opposite directions. 

    The upper second stage gear drives the tubular lower rotor shaft clockwise (viewed from above) and the lower second stage gear drives the tubular upper rotor shaft counterclockwise within the hollow center of the lower rotor shaft. Total gear reduction from engine output to rotor shafts is 9.83 to 1. The transmission gears, shafts, and bearings are contained in a housing, which consists of three major sections. The main transmission housing, in the center, contains the first and second stage gearing. The upper conical bell housing section contains the main rotor thrust bearing. This thrust bearing is of the tapered roller type. 

    The lower section constitutes a transmission oil sump and a housing for the displacement type oil pump. Oil is fed at a nominal 60 psig to all gears and bearings in the transmission as well as to the rotor shaft bearings. Scavenging is by gravity feed. Oil used in the transmission is Specification MIL- L-23699. The oil fill cap is on the upper left side of the transmission housing. The transmission oil sight gage is located on the lower left side of the transmission. Sockets are provided in the lower transmission casting to receive the landing gear struts and the aft fuselage members. The bomb shackles also are supported by this casting. Torque is imparted to the upper and lower rotor shafts by the lower and upper second stage gears, respectively, through splines. 

    Upper rotor shaft thrust is transmitted to the lower rotor shaft by two thrust bearings at its upper end. Thus the total rotor thrust is transmitted to the upper section of the transmission housing through the main thrust bearing mentioned above. The splined couplings between the second stage gears and the rotor shafts are capable of limited axial movement, eliminating the transfer of rotor thrust loads to the gear hubs, and consequently to the mesh. 

    The accessory drive gearing is incorporated in the upper transmission housing. This gearing transmits power from the rotor drive system to the airborne generator and the four-axis rotary servo actuator. A bevel gear, keyed to the hub of the lower rotor shaft second stage gear, provides means for power takeoff. Rotor shaft speed is stepped up to 1736 rpm for the input to the servo actuator, and to 8000 rpm for the airborne generator. An overrunning clutch in the accessory drive gearing to the servo actuator permits the application of auxiliary prime mover power to the actuator for test purposes.

DRONE ELECTRICAL POWER SUPPLY SYSTEM

Electrical power for all the airborne electrical circuits is provided by an airborne generator, which is driven, through gearing, by the rotor drive system. The generator provides the following nominal voltages

115 volts, 400 cps, phase A, B, C.

22 volts, 400 cps, phase 1, 2, 3.

26 volts, 400 cps, phase A, phase A-90 degrees.

A three-phase rectifier circuit in the relay assembly, operating from the 22-volt 400 cps supply, provides unfiltered 28-vdc power to operate relays, solenoids, and other equipment. The 22-volt generator winding also supplies power to two three-phase rectifier circuits in the electronic control amplifier, one of which provides an unregulated 28-vdc supply for internal use within the unit.

The second rectifier circuit is used in conjunction with the 28-vdc supply in a voltage doubler circuit to produce a 39-vdc supply for transistor biases.

Power is transmitted to the various units of the system by means of wiring harnesses and cables.

LANDING GEAR

The landing gear consists of four tubular struts and two tubular skids. Fittings at the ends of each skid include provisions for clamping the drone to the deck in the hangar, and for the installation of restraint rollers for deck handling.

An electrical switch is mounted on each skid. These switches are actuated when the drone makes or breaks contact with the deck. When the drone is on the deck, the switches serve to short out follow-up washout circuits in the AFC set. When the drone is airborne, the switches serve to open the engine-off command circuit so that the drone engine cannot be shut off remotely while the drone is airborne. The control lock and under voltage cutout circuits also are interlocked through these switches.

AFC SET AND RECEIVING SET

The components of the AFC set and receiving set are listed in figure 3. The servo actuator is mounted on the forward side of the accessory drive section of the transmission, and receives its mechanical input power from the accessory drive gearing. The airborne generator is mounted on the after side of the generator support housing assembly, and is driven by the accessory drive gearing. The decoder is bracket mounted on the upper tubes of the aft fuselage assembly. The control amplifier, radio receiver, and relay assembly are mounted on the after side of the avionic panel, which is supported by the tubing of the a