The Curtiss Commando has been recently introduced into the field of post-war commercial transports. Originally designed as a commercial transport in 1937, it was redesigned as a military cargo and troop transport in 1940.
The expediences of military operations exaggerated many of the minor shortcomings of prewar commercial transports. Accessibility and simplicity of equipment became prominent factors since they are of utmost military value. After the incorporation of over a thousand design changes, the C-46 has approached mechanical perfection.
This backlog of experience gained with the C-46 in war operations was transferred directly into the design of the CW-20E, and greatly simplified the task of designing maximum efficiency into this commercial transport.
The Wright Cyclone 18 engines equip the plane to meet demands for large payloads, great reliability and economical operating costs, its performance being particularly outstanding on flights up to 700 miles a range that accounts for over 90% of domestic air travel. With a top speed of 286 mph, and cruising at 251 mph at 14,000 ft altitude, the public's demands for speed are more than met. By efficient design, direct flying costs have been lowered to 9¢ per ton mile.
Modifications of the cockpit provide the highest degree of utility and comfort for the pilot. The pilot's pedestal has been modified by removing several controls, relocating others, and streamlining the exterior into a more attractive and compact configuration. This change improved access to the pilot's seat, and the reshaped and colored control knobs facilitate instantaneous distinction between controls, resulting in greater ease of operation.
The landing gear selector was modified for increased safety when in the ground. Safeguard against accidental retraction on the ground is now achieved by the modified selector knob which requires rotation as well as depression to release the selector from the down position. For added protection a device is employed which makes it impossible to retract the landing gear while it is supporting the weight of the airplane. This is achieved by the incorporating of a mechanism on each main landing gear strut which will not permit the secondary mechanical downlock to be released by the landing gear selector when the main struts are compressed beyond 1 in from their full extended position.
The instrument panel was rearranged to furnish an identical subpanel of flight instruments for both pilot and co-pilot. The automatic pilot and engine instruments are located in the upper center of the area of the main panel for availability to both pilots. As an aid to serviceability and maintenance, the flight instrument subpanels are interchangeable.
Night and instrument flying are made easier by reflection-free cockpit illumination. General lighting for the pilot's compartment is provided by one 21-cp light. The instrument panel is illuminated by four fluorescent adjustable lights rheostatically controlled. The instruments are also illuminated by five incandescent lights mounted around the upper extremities of the instrument panel and are rheostatically controlled. An incandescent map reading light is installed on each side of the cockpit, with the controlling rheostat adjacent. The side panel is illuminated by a light mounted on each side wall and the pedestal is lighted by two pin point spot lights in the overhead panel. Four hooded lights illuminate the overhead panel itself.
All radio controls, except the sensitivity control, are located in a consolidated control panel just above the center of the windshield, forward of the pilot's shoulder, and are accessible for operation to both the pilot and co-pilot. With the panel located in this position, no backward reaching is necessary to adjust the radio controls. The panel is accessible for maintenance by operation of slide fasteners requiring no tools, and a quick-removable hinge allows complete removal of the entire panel for replacement. Dual sensitivity controls are installed on either side of the pilot's pedestal for convenience of operation.
In order to achieve ideal visibility characteristics for the pilot, the entire nose and cockpit of the CW-20E was redesigned to incorporate a flat glass windshield with negligible refractive and reflective errors. At the same time, the field of vision was increased.
Vision is provided for the pilots through the two-panel forward windshield, one oblique window in the corner, and one window on each side. All these windshield panels are flat plate glass laminated with vinyl plastic. The two panels of the front windshields are double with a space between to permit heated air to be circulated for defrosting and de-icing. Inside panels are hinged and open inward for cleaning.
An average pilot in flight position has an upward angle of vision out the front windshield panel of 16°, and a downward angle of vision of 15.5°, measured from a horizontal line at eye level.
Distance from the normal position of the pilot's eyes has been reduced from 30 in to 18 in. This increases the visibility by increasing the angle between upper and lower limits of vision. The pilot may, by leaning forward, increase this angle still more. This installation is of bird-proof construction.
Information indicates that the greatest protection against heavy birds is a thick vinyl. The glass in such an installation shatters but this heavy vinyl acts as a net and prevents a bird from entering the cockpit even though it comes through the heavy forward piece of glass. The rear panel is attached to the airplane structure along the entire length of its upper and lower edges. The upper edge support consists of a continuous hinge and the lower edge support consists of a quick-operating latch arrangement which furnished a maximum of supported length.
Both panels of the windshield are so constructed that they are removable as a unit with the frame. This means that the forward panel can be sealed in its frame at the factory and tested before installation in an airplane. This outer unit is installed with a soft rubber gasket to eliminate distortion which might be caused by the panel being tightened down on a hard surface.
A clear-vision panel which may be opened in flight was provided in order to make it possible to obtain an unobstructed view ahead under extreme rainy conditions. This panel also makes it possible to clean the front of the windshield manually in flight.
A completely new mechanism has been devised for opening and closing the side windows. This mechanism operates smoothly and freely and provides an outward movement at the end of the closing travel that effectively seals these windows against leakage.
The new heated-air windshield is believed to be the first windshield de-icing system designed that will prevent or entirely remove ice under the worst conditions encountered in flight.
After considerable research and design studies, a trial installation of the windshield heating system was installed in a C-46 for flight testing. The plane was equipped with a water spray system so that raw water could be sprayed on the windshield in such a manner as to simulate the most severe icing conditions. Ice formed by this system was known as "glaze" ice which is the most tenacious form of ice developed.
The hot air blower system passes 1000 Btus per hr per sq ft of windshield area, and 550 lb of heated air per hr to the windshield. The windshield de-icing blower, located in the cockpit assists the main heating system blower in supplying air to the windshield.
For the flight test, the airplane was flown at an altitude of 20,000 ft with the test being made at two different cruising speeds, the first at 175 mph, and the second at 248 mph. The results of the test showed that the airspeed had only a negligible effect on the effectiveness of the de-icing system. At this altitude the outside air temperature was +15°F. When the temperature of the windshield had stabilized, the water spray system was turned on so that water was sprayed evenly over the surface of the windshield. When a 1/16 in of ice had formed, the spray was turned off. The windshield de-icing heat valve was now opened so that the heated air could pass between the two layers of the windshield glass. Withing 2 min, over 70% of the ice had been removed and the remaining ice was thin and clear so that visibility was unobstructed. Within another few minutes all the ice was completely removed and the windshield was dry.
In an attempt to eliminate complex and duplicating mechanisms, Curtiss engineers tackled the problem of eliminating the troublesome hydraulic boost control system that had been found necessary on all large aircraft, and replacing it with a simple, foolproof mechanical system that would still retain minimum pilot control forces. This program was guided by former experience, especially on the Curtiss C-76 transport, and by laboratory research.
As a result of the boost elimination program, Curtiss-Wright has designed and flight tested a complete set of flight controls that have exceeded all original expectations.
The aileron is an aerodynamically balanced Frise type incorporating a combination trim and balance tab. The aileron is of all-metal construction and metal covered, incorporating a full span spar and one partial span spar for the tab hinge. Spars and ribs are of flanged sheet metal with lightening holes. Each aileron is attached to the outer panel by four anti-friction ball bearings. Differential motion is obtained through a bell crank and push-rod actuated by the aileron control cables. The aileron is removable by disconnecting the controls and removing the hinge bolts.
Elevators and rudder are metal covered and of all-metal aluminum alloy construction, comprising two longitudinal beams and stamped sheet metal ribs. Elevators are attached tot he stabilizer by means of five anti-friction ball bearing hinges on each side. The rudder is attached to the fin by means of six anti-friction ball bearing hinges. Elevators are controlled through a steel torque tube which bolts to the inboard ends of the surfaces and are actuated by means of horns, to which the control cables are connected. The elevator torque tube is continuous between the elevators and is supported on the two hinges that are within the fuselage.
Each control surface is provided with a trim tab located in the trailing edge of the surface. The trim tabs are all mechanically actuated; trim adjustment is controlled from the cockpit. They are all-metal construction and each is attached to the control surface by anti-friction ball bearings. Trim tabs are connected to their actuators by push-pull rods.
Spring control tabs are provided in the trailing edge of each elevator and on the rudder. These tabs reduce the pilot force necessary to actuate the control surfaces, and maintain "feel." This is accomplished by a linkage system, proportioning the pilot force to the tab as well as to the main control surface. Included in the linkage system is a preloaded spring whose preload must be exceeded in order to move the tab with respect to the control surface. These tabs are attached to the surface by anti-friction ball bearings.
Through the use of spring tabs, aerodynamic balance, and a new Curtiss-Wright mechanism termed the "Vee Tab," the CW-20E Commando has not only eliminated all necessity for hydraulic boost controls but at the same time achieved finely balanced and coordinated flight controls, and greatly increased the CG loading range.
With the knowledge that many airplane accidents have been caused by failure of pilots or ground crews to remove or release control surfaces gust locks prior to flight, Curtiss engineers attacked the problem of designing a foolproof gust lock that would eliminate this hazard to flying. A system of linkages has finally been worked out that locks the control surfaces against any loads applied on the surface, but may be unlocked simply by a small movement of the flight controls.
The gust lock problem was fortunately simplified by the use of spring tabs in the control system. The locking mechanism consists essentially of a three-link dead-center system similar to a trunk latch. The travel of the spring tab is initially independent of the control surface and, therefore, provides the necessary motion required to automatically unlock the system through motion of the cockpit flight controls. Should the pilot inadvertently take off without having released the gust lock, his first movement of the flight controls will automatically do so.
In order to provide ease of accessibility, the battery was relocated in the fuselage lower surface at a position just forward of the heater access door and mounted on an elevator type carriage. This feature was originally designed in the early Condor transports and, with refinements, again adapted to the CW-20E Commando. The carriage is raised or lowered by an attached, flush, jackscrew hand crank, and is counterbalanced so that it requires little effort for raising or lowering.
The main cabin is ventilated by heated or fresh air introduced through outlets in the heated air ducts which extend the length of the cabin at either edge of the floor, and mainly exhaust upward through the cabin ceiling duct and fluorescent light fixture. Partial exhausting of the cabin air is forward through the galley section and rearward through the dressing rooms, for the prevention of any odors, originating in these sections, from being conducted to the main cabin.
Since the postwar era will demand the ultimate in comfort, the passenger compartment is richly appointed, with a fully automatic temperature control and finely developed system of ventilation.
The program for developing the interior cabin design included painstaking research and innumerable conferences with transport operators, pilots, passenger service personnel, and engineers. From the data gathered, Curtiss engineers designed the passengers' facilities to provide the utmost in comfort and convenience and yet, at the same time, provide a fully utilitarian design, simply constructed, easily serviced, and requiring little maintenance.
For the first time in commercial transports, fluorescent indirect lighting is employed to illuminate the passenger compartment. This revolutionary system of lighting, supplemented by the reading lights, provides over four times greater candle-power at reading level than present day transports. A soft uniform light is obtained throughout the length and breadth of the cabin without annoying reflections, spotty areas of illumination, or shadows. An additional advantage obtained by the use of fluorescent lights is a considerable savings in weight over an incandescent installation, at the same time achieving a more simple installation with easier servicing. The protective reflector shield is cut into short segments with hinges on one edge and fastened in place on the other edge with two finger-operated slide fasteners. With the reflector hinge downward the complete lighting fixture for two units is exposed for easy servicing and replacement of tubes. The automatic self-starting units are removed by a quarter turn of the unit in its socket.
For the relief of air sickness and individual temperature control, each passenger is provided with an individual fresh air grill, in the overhead rack, with variable volume control.
The dual passenger lounging chairs are deeply upholstered with foam sponge rubber, leather edged, and are finger tip controlled for reclined position. A spacing of 41 in between seats permits ample leg room and does not crowd the other passenger when the seat is reclined. All seats are interchangeable, reversible, and easily removed simply by the operation of one finger-operated lever that releases the seat from the floor. The seat is held to the floor by means of six clevis plates, on the pedestal frame, that slide under six studs mounted on the floor. The quick release consists of a plunger rod that drops into a shear fitting in the seat fitting and cabin floor structure. To remove the seat, the seat release plunger is turned a quarter turn and pulled upward. The entire seat may then be slid toward the aisle and removed as a unit from the airplane. By this means, any number of seats may be easily and quickly removed to provide space for cargo, in emergencies, or to facilitate cleaning. The reversible feature permits rapid conversion of any set of seats into a club car effect with seats facing each other.
Sixteen large windows, 16 in x 14 in, are conveniently located so as to furnish each passenger with a panoramic view. The plastic windows are easily removable from the outside for replacement. The wide window ledges are tilted outward to provide a shelf for cigarettes, handbags, writing material, and other small items.
Two twin lavatories are installed in the aft end of the fuselage. The wash basin and the adjacent table top is made of buffed aluminum. An attractive lounge type toilet is so designed as to also serve as a lounge chair when not in use as a toilet.
The CW-20E Commando galley is so designed as to provide facilities for almost any type of food service desired. Space is provided for a service room or "galley" between the pilot's compartment and the passengers' cabin. The floor area is 30.6 sq ft and the volume of the galley compartment space is 186 cu ft.
All cabinets are made of stainless steel and provided with adequate drainage for corrosion resistance. Cabinets are designed in sections to they may be easily removed from the airplane. Each section may be removed individually simply by the removal of from four to six attaching studs.
Alternating and direct current, and an outlet for a 1500W hot plate or baby bottle warmer, are furnished in the galley. Illumination is provided by indirect fluorescent lighting and one incandescent light, both controlled by switches in the galley.
The power plant is designed as a single unit or "power egg." The unit is complete within itself, including propeller, engine, engine mount, engine accessories, oil tank, firewall, and cowling.
The two-zone installation incorporates a stainless steel diaphragm ring that separates the nose section from the accessory compartment and assists in directing the cooling air exit flow.
Engine mounts are chrome molybdenum steel tubing with a steel tube engine mount ring. The mount is bolted to the nacelles through four fittings, incorporating spherical contact surfaces, located at the ends of the nacelle longerons. The engine is supported from the mount ring through an anti-vibration suspension consisting of a series of rubber shock mounts.
Removal of the "power egg" does not require re-rigging of control cables or draining of airplane fuel, oil, or hydraulic systems. Engine sections are interchangeable left and right.
Firewalls are constructed of a single sheet of stainless steel reinforced to support the oil tank and other power plant accessories which are supported from it.
Engine cowling framework is supported from the engine mount structure and consists of a nose ring supported by four longitudinal box-section beams. Each beam is attached directly to the engine mount at the main engine mount attachment fittings and is supported at its approximate midspan by struts from the engine mount ring. The nose ring is fastened to the beam by means of bolts.
The engine cowling is divided into eight panels which are hinged to the cowl beams, and secured by means of toggle latches which may be operated by hand without the use of tools. In the engine section of the cowling these panels extend from the nose ring to the cowl flap support ring and comprise the two side panels between the cowling beams, the top panel with the engine charge air intake duct, and the bottom panel with the oil cooling air intake duct. In the accessory section, these panels extend from the diaphragm support ring to the firewall and comprise the two side panels, the top panel between the cowling beams, and the bottom panel.
Nose section cowling incorporates an air duct and an air filter. The nose ring is removable without removing the propeller. In the upper portion of the leading edge is incorporated the ramming air intake. An intake to supply air for the oil coolers and for accessory compartment ventilation is incorporated in the lower portion of the cowl leading edge. The charge air intake duct is integral with the forward top cowling panel. Oil cooler ducts are integral with the forward bottom cowling panels.
A cooling air diffuser is provided as an inner lining of the nose cowl. Cooling air flow through the nose cowling is regulated by cowl flaps at its trailing edge. These flaps are of the airfoil section type and are so hinged that, when open, a slot is formed at the leading edge of each flap, directing the airflow over the cowl to increase the cooling and decrease the drag. Flaps are actuated by an electrically-driven screw mechanism controlled from switches in the pilot's compartment. A cowl flap position indicator is provided in the pilot's compartment. Cowl flaps are capable of being opened sufficiently to allow servicing of installations through the opening formed.
A fixed exit grill is provided at the accessory cowl trailing edge for accessory section ventilation.
The engine is cooled by air flow through the engine cowling. A duct from the oil cooler duct in the lower section of the cowling provides cooling air for the generator which is passed through a filter.
Engine oil is cooled by means of a 17-in diameter, 9-in long, .210-tube, anti-congealing radiator, supported from the engine mount and located in the lower cowl duct. Oil temperature is controlled thermostatically by means of an electrically operated, automatically controlled air exit flap with a manual override switch.
Rammed charge air is taken in through an opening in the upper section of the cowling leading edge, and passes through a duct in the upper portion of the cowling.
Heated air is supplied to the fuel metering control by passing air over a plate-type heat exchanger located in the intake air duct elbow. Engine exhaust gases are passed through this exchanger, only when heated air is desired. The flow of exhaust gases is controlled from the pilot's compartment. When heated air is not desired the exchanger is bypassed and the exhaust gases are discharged through the manifold. The heat exchanger is designed for a minimum loss of ram pressure with or without heat. The system is designed to provide a heat rise of 100°F, at 30°F ambient temperature and 75% METO power.
A separate oil tank is provided for each engine. The oil tank is of welded aluminum alloy construction and mounted just forward of the firewall in the engine section. The oil radiator is automatically controlled and automatically protected against congealing and surges due to cold starts. The radiator is mounted on rubber shock mounts in the lower portion of the engine section. An oil dilution system to facilitate starting in cold weather is installed, including a hopper in th oil tank, a solenoid valve, and connection to the fuel pressure line. The entire lubrication system, including the oil tank and lines forward of the firewall, is removed with the engine assembly. Disconnect fittings are provided at the firewall for all lubrication system lines or controls passing through the firewall.
The main fuel system comprises four fuel tanks, two located in each outer wing panel, together with the necessary pumps, strainers, lines and fittings. Total fuel capacity is 1056 gal, with space provisions for a third 172-gal fuel tank aft of the second tank, in each outer panel.
Fuel tanks are constructed of welded aluminum alloy. They are supported in padded cutouts in specially designed wing ribs. The support allows no appreciable movement of the tanks, or flexing and bending to be transmitted to the tanks. An expansion space of 8 gal for each tank is provided to prevent filling this space when the plane is on the ground in the three-point attitude. Fuel tank filler caps are flush with the wing surface and incorporate pressure relief valves. A hopper and drain at the filler neck opening prevents fuel spilled at the filler opening from entering the wing, fuselage, or nacelle. Tank compartments are vented and drained to prevent fire explosion. Tank outlets supplying fuel to the system are baffled to prevent uncovering the fuel outlet and interruption of the fuel supply while the airplane is in unusual maneuvers or side slip. Each tank is equipped with a fuel quantity transmitter of the float type. Right and left front tank transmitters are connected to one dual indicator on the instrument panel. The rear tank transmitters are connected in a similar manner.
The crossfeed line connects to each system on the pressure side of the booster pump, permitting fuel to be pumped from one system to the other system. The crossfeed valves are cable controlled from the pilot's compartment.
Each engine has a separate fuel system furnished by the two fuel tanks on that side, each system operating as a self-contained unit. The fuel may be drawn from either one of the two tanks through a selector valve controlled from pilot's compartment. An electrically driven booster pump is installed to provide for emergency, takeoff, or high-altitude operation. These pumps are controlled by switches in the pilot's compartment.
This Design Analysis article was originally published in the March, 1945, issue of Industrial Aviation, vol 2, no 3, pp 7-8, 10, 12-14, 16, 18-23.
The original article includes a thumbnail portrait of the author, 4 photos, 12 detail drawings and diagrams, and 3 data tables, plus a ledger-sized foldout with a color phantom rendering of the airplane in civilian passenger-transport configuration.
Note: the copy from which this is taken is missing the foldout. The phantom rendering used to develop the wallpapers was taken from Flying magazine.
Photos are not credited.