The Lockheed P-38 Lightning is unique among fighter aircraft, due principally to its versatility. It is used as a high-, low- and medium-altitude fighter, interceptor, bomber escort, light bomber, skip bomber, dive bomber, attack bomber and strafer, supply dropping plane, night fighter, "piggy-back" fighter-trainer, tank buster, smoke screen layer, and as a photo-reconnaissance plane. It has been tried as a torpedo plane and a glider tow, and before this war is over it will be put to other important uses which for the moment must remain undisclosed. Distinctive in appearance, P-38 is among the largest and heaviest fighter planes. It is one of the world's fastest also and at its best altitude, it outspeeds any other production fighter. Its single engine speed is in excess of 300 mph, and its service ceiling, where it still climbs 100 fpm is considerably in excess of 40,000 ft.
P-38, first conceived in 1937, has evolved through 16 major modifications to the P-38J, latest of the series to actually be delivered, although others are in process of engineering, experimental work or testing. Basically, it is an all-metal, midwing, twin-engined single seater having twin boom nacelles and fuselage of semi-monocoque, stressed skin type.
The wing is full cantilever, stressed skin type and consists of a center section, two outer panels, and wing tips. To the wing are attached the two engine nacelle and the gondola type fuselage or "pod." Two Allison engines are mounted outboard in separate nacelles. The engine nacelles fair back into the booms which support the tail unit, their streamlining broken only by the coolant radiators located approximately halfway aft from the wing centerline. There is no orthodox fuselage, the gondola being at the trailing edge of the wing at the plane of symmetry and extending forward only. Separate compartments are provided for armament, (all guns are concentrated in the nose), cockpit, retractable nose wheel, radio, hydraulics and fuel. Since the plane sits level on its tricycle gear, access over the wing is provided by a retractable ladder mounted in the aft end of the fuselage. Entrance to the cockpit is by hinged canopy used in con- junction with windows on both sides of the cockpit enclosure.
The body group, consisting of the center section, fuselage and forward booms, is jig-mated, riveted and bolted, and as such is considered an irreplaceable unit. The aft booms are also jig-mated and bolted but may be replaced.
The fuselage is an all-metal, semi-monocoque gondola-type structure, framed primarily with bulkheads hydro-press stamped from 24ST sheet with transverse partitions between the different compartments which provide additional strength. Longitudinal stringers are of bent up sheet aluminum alloy, the skin varying from .051 to .025 thickness. Skin sheets are butt-jointed and flush riveted. Armament compartment doors are built up of a smooth outer skin and a formed inner-skin strengthener. Both are stamped from .040 material, and are riveted and spot- welded. The door of the nose wheel well is similar in construction except that .032 Alclad is used for the outer skin and .040 for the inner skin. There is a firewall between the armament compartment and the pilot's station additionally reinforced by armor plate.
The lower section of the fuselage contains the nose gear, a compartment for receiving ejected cannon shell cases and links, major units of the hydraulic pressure control system, and control valves for flaps and landing gear. Fuel valves, strainers, booster pumps and the engine and surface-control cables also are located here.
The upper section contains the armament, cockpit and radio compartments. Access to the armament compartment is through two hinged panels secured by Dzus fasteners. The aft section of the fuselage contains the hydraulic reservoir, flap motor drive and mounting ladder. The ladder is pivoted and spring loaded so that it retracts into the tail cone; a hand hole is provided in the left side of the fuselage for assistance in climbing the ladder.
The windshield is composed of three separate glass panels. The front panel is bullet-proof, being made of five layers of glass set in vinyl plastic. (This differs from an earlier arrangement of a curved Plexiglas enclosure with a flat glass shield mounted inside the cabin before the pilot.) Side panels are the same but with two layers of glass, are mounted in synthetic sponge rubber in an aluminum alloy frame and held in place with aluminum strips.
The side windows in the cockpit enclosure lit in slots between the front shear beam and the main beam, raised and lowered by a hand crank and locked in the raised position by a ratchet and pawl. The cockpit hatch is a molded plastic panel fitted to a metal frame. The hatch opens aft on a hinge and incorporated in the hinge is a telescope stop. A pin and cable assembly, controlled from in or outside the hatch, locks the hatch in the down position. A red-painted emergency handle is provided in the cockpit. If hatch is released by this handle in flight it raises about 2". Air pressure will then tear the hatch away leaving free emergency egress for the pilot.
The aft canopy is made up of two plastic panels ¼" molded sheet set in metal frames and secured by Airloc fasteners for quick removal. These act as a cover for the radio compartment.
On the underside of the fuselage is the nose gear compartment and door. The gear retracts directly aft and upward in its fully extended position and the door is automatically operated by a hydraulic cylinder connected directly to the aft end of the door. A hydraulic latching device holds the front end of the door tightly closed.
The forward booms extend from the engine firewall to just forward of the coolant radiators and are jig-mated to the center section. The main landing gear and turbosupercharger installations are contained in these sections. Stressed skin construction is used including the inside skin of the wheel wells, except for that portion of the upper surface which is exposed to the supercharger heat. This area is fabricated from .018 stainless steel.
Internal construction consists of stamped aluminum bulkheads and bulb angle stringers, the bulkheads spaced about 15" on centers. As in the after booms the skin thickness is .032". Forward booms are somewhat elliptical also, and there is no abrupt change.
The main landing gear retracts aft into the wheel wells with a pendulum motion. Two doors, hinged to the lower channels of each forward boom operate automatically with the landing gear movement. A hydraulic cylinder located in the aft end of the wheel well operates the front and rear door carriage through a linkage of cables and rods.
The aft booms are of semi-monocoque stressed skin type, and extend from the forward boom to the empennage boom. Carrying the coolant radiator structures on either side of each boom. Attachment of the aft booms at both ends is done with screws and stop nuts through the skin and webs, and by bolts through two fittings forged from 14ST aluminum that mate with fittings on the forward boom channels. Coolant radiator frames are supported by brackets attached to the formers. A baggage compartment is located in the right-hand aft boom.
Each end of the compartment is closed off by an aluminum alloy bulkhead, the forward bulkheads retaining the aircraft data case. A similar compartment for battery stowage is in the left-hand aft boom. Access to the interior of the aft booms is through manholes on the underside. The aft booms also provide space for large oxygen supply bottles.
The pilot compartment is situated immediately forward on the main wing beam and is roomy and well arranged. Pilot has full 360° visibility through bullet-proof glass panels. Control is of the wheel type with cables carried from the three-spoke, two-thirds wheel to the right of the cockpit through an inverted L-shaped control column. The cannon and machine gun trigger button is located near the pilot's right thumb on the wheel and a radio switch adjacent to it. The master armament switch is mounted on the forward side of the horizontal member of the control column. Engine and propeller controls are grouped at the left while flight and engine instruments and most switches are located on the main panel directly before the pilot. A map case and first aid kit are carried behind the seat on the right side and a flare pistol assembly is on the left. The seat is molded of plastic, and both seat and back are protected by 3/8" face hardened steel armor plate. Regulation seat belt and quick-acting shoulder harness are provided and a pilot relief tube is at the left edge of the seat. Silhouette armor plate is placed between the cockpit and the radio compartment immediately aft.
The wing group consists of the center section, two outer panels and the extreme wing tips. In the center section, a main beam and front and rear shear beams are the main structural members.
The main beam is located at 35% chord and is of double-web construction. It is built up of top and bottom cap strips of 24ST aluminum extrusions and double webs which are .064" in thickness in the cockpit section and .040 outboard. The forward shear beam does not continue through the cockpit section and does not extend into the outer wing panels. It serves as a strengthening unit between the cockpit structure and the outboard ends of the center section on each side. The extruded cap strip and web construction is used on all wing beams with the exception of the inboard front web of of the main beam in the center section. Here, a built up truss section is used for added strength. Longitudinal ribs are used in the center section except between stations 17 and 68 between the main and rear shear beams, and between stations 18½ and 79½ between the main and forward shear beams. (In the Lockheed production setup, stations are 1" apart from the plane's centerline). In the aft section where the main tank is installed, rib strength is replaced by a corrugated inner skin of .040 Alclad, and in the forward section where the reserve tank is installed, hat section formers are used to give the necessary support.
Outer panels, beams, main and rear shear, join the corresponding beams in the center section. While the main beam continues directly outboard, the rear shear beam runs parallel to the trailing edge of the wing to a point just outboard of the inner end of the aileron. Here it is spliced to a lighter outboard section and the load at the splice is carried through the surface structure and the ribs. The main beam in the outer panel is of single web construction with extruded top and bottom cap strips. The rear shear beam employs cap strips of bent-up sheet angles formed from 24ST with a web of .051 sheet strengthened by irregularly spaced angle stiffeners. The main beam cap strips are formed of double reinforced L-shaped extrusions which reduce uniformly in section as they go outboard. Roughly the outer half of these caps is of sheet angles rather than extrusions. At the inboard end of the main beam in these outer panels all vertical shear is carried through a diagonal hat section to the top beam fitting, the bottom fitting securing the lower side of the wing.
Outer skin of the wing is in effect a double layer composed of a corrugated inner skin and a butt-jointed, flush riveted outer covering. Corrugations in the center section are of .064 24ST and the skin of 0.40; on the outer panels, corrugations are of .032 material on the lower surface and .064 on the upper. Outer panel skin varies in thickness from .020 to .040.
The wing panels are skin stressed only between beams. The trailing edge section is composed of ribs and skin on the upper surface with intercostal stiffeners between the ribs. In the latest model, the leading wing edge contains shells, and is composed of chord-wise corrugations and smooth skin, with few ribs. Outer skin is attached to 14 ribs in the outer panel, the ribs being formed from 24ST sheet varying in thickness for .025 to .025".
Outer wing panels are solidly joined to the center section. The main beam fitting is a multiple-lug, pin-jointed connection at top and bottom of the beam. The center section fitting is a 14ST aluminum forging, the outer panel fitting is forged from 4130 steel. The pins are also of 4130 steel, the upper 11/16" in diameter and the lower 5/8". The rear shear beam is joined by steel fittings bolted together at top and bottom. For further strength, 9 upper and 9 lower bathtub fittings are provided between the two beams. These are 14ST aluminum forgings and are carried well back into the corrugations of the inner skin for strength. They are bolted together with 7 /16" high-strength, internal wrenching, tension bolts on the lower side and similar 5/16" bolts on the upper fittings.
Wing flaps are Lockheed-Fowler type which roll aft and down from the trailing edge on tracks beneath the wing, extending the actual area of the wing as well as increasing lift. Relatively slow landings of highly wing-loaded aircraft are made possible because of the high lift coefficients. With the flap in a partly extended position, a large increase in lift is provided with a minimum increase in drag. This makes possible the use of these flaps for take-off as well as for a maneuvering flap in combat. The P-38 is the first and only fighter airplane equipped with this type of flap.
Flaps, divided into four panels, are interconnected and operate together. They are installed in the lower wing surface between wing stations 8 and 77¼ in the center section, and between 118¾ and 180 in the outer panels. At each end they are attached to carriages that roll in tracks built into the wing structure. The carriages are linked by cables to push-pull tubes, traveling in roller brackets on the rear face of the rear shear beam. Push-pull tubes are actuated by long screws driven by a hydraulic motor housed under upper rear fuselage section.
All movable control surfaces are similar in construction. They are built up of ribs, stringers, and outer skin with internal spars of sheet metal angle and stamped web construction. Unlike the wings these surfaces have no corrugated inner skin. Outer skin varies in thickness from .020" to .032" on the long horizontal stabilizer. All control surfaces are actuated by cables which run outboard from the cockpit through the center section, the tail surface cables being carried aft on pulleys.
The twin-engined arrangement of the P-38 was specified in the original design in 1937. At that time no single engine available would provide the horsepower needed for the performance the Army wanted. Although engines have been increased in power since then, the dual installation of the P-38 still makes it the most powerful fighter in the world. In addition, combat reports prove conclusively that the double-powered fighter has many advantages, the greatest of which is safety. Lightnings have come home from combat engagements on one engine so frequently that they are known as "the round-trip ticket." It has been shown too that enemy fire most frequently enters a fighter plane through the fuselage aft of the pilot station. Since the P-38 has no fuselage aft of the wing, enemy bullets pass harmlessly through the open section between booms or enter the booms themselves where they do little damage.
Locating the engines remote from the pilot and armament allowed for wide latitude in power plant installation and provided widely spaced main wheel wells for maximum tread and landing safety. This arrangement also removes from the vicinity of the pilot the hot Prestone, hot oil and much of the fuel. Such inflammables are in separate engine nacelles on each wing where a power plant fire may be dealt with by side-slipping, or isolated and extinguished. It also lessens chance of explosion in combat.
One of the most important combat advantages of the twin-engine design is that it permits the concentration of all firepower in the nose. The trajectories of the fixed guns are parallel with the pilot's gunsight and maximum firepower is possible at any point directly ahead of the plane up to the range of the guns. This is not true in wing installations which feature a "cone of fire" for which there is only one optimum range.
Engines are mounted on forged aluminum alloy of heavy construction and of triangular shape, and bolted to fittings at the forward corners of the support bay. The bay is also of forged aluminum alloy and is joined by two large machine bolts to heavy fittings on the forward boom wheel well channels. Both the support bay and the longitudinal trusses are supported by tubular diagonal hangers bolted at their upper ends to fittings on the cross members of the forward boom.
The P-38J is powered by two, 12 cylinder, V-1710-F2, liquid-cooled Allison engines. The left propeller rotates clockwise and the right propeller counterclockwise, viewed from the front of the airplane. This counter-rotation feature results in the elimination of torque with resultant improvement in flight and ground handling characteristics. The engines are sea-level types and are constructed for adaptation of exhaust turbosupercharging to maintain full sea-level ratings at altitude. Each engine has a self-contained, single-stage blower, with a 8.10:1 blower-gear ratio driving a 9½"-diameter impeller, located in the accessory housing section in the forward boom; and an auxiliary, General Electric, Type B-33, exhaust-driven turbo supercharger, mounted in the forward boom.
The engines develop 40% more horsepower than the original P-38 installation. They have a bore of 5.50", a stroke of 6" and a piston displacement of 1710 cu in. Normal rated speed is 2600 rpm or 3000 for take-off or military needs. Their rated bhp is 1100 normal at 2600 rpm and 1600 for take-off or combat at 3000 rpm. The propeller reduction gear ratio is 2:1 and the propellers turn in opposite direction to the crankshaft rotation of the engines. Each engine weighs 1350 lb complete, is 85-29/32" long, 29-9/32" wide and 36 24" [sic] high. Each engine is mounted by eight bolts .4375" in diameter and spaced 18-3/8" apart.
Magnetos are Scintilla double-fixed timing type and turn at 1.5 times the crankshaft speed. AC or Champion spark plugs are specified. Carburetors are Bendix-Stromberg and 100 or 100-plus octane fuel of 130 or 140 grade is used. Fuel pressure is 16-18 psi at 2200 rpm. Maximum oil consumption is approximately 15.5 qts per hr at rated power at 2600 rpm. Minimum temperature allowable for take-off or flight is 100°F, and maximum 194°. Engine starters are of the electric-inertia type although provision is made for emergency hand-inertia starting and a crank is mounted in the wheel well.
Propellers are Curtiss, three-bladed, full-feathering, controllable pitch type. They counter-rotate out from the top of the propeller swing. On the P-38J, the blades are of a new type which give much better performance at extremely high altitudes. Props are 11' 6"-in diameter and are geared down 2:1 from engine rpm.
Engines are cooled by ethylene glycol with a separate system for each engine. Coolant radiators are located midway of the aft booms and the temperature of the coolant is regulated by varying the flow of air through the radiators. Hydraulic exit flaps for the radiators are automatically controlled through a temperature-reactant four-way valve. An expansion tank is provided and connected to the coolant pump inlet.
An independent pressure lubrication system provides oil for each engine. AiResearch temperature regulators, two for each engine, are mounted under each engine. An automatic exit flap and electrical actuator motor control the flow of air through the regulator and a bi-metal line thermostat is installed between the regulator and the tank to control the position of the exit flap. A surge valve limits Slavanger pump pressures to 45 psi or less, and an oil dilution system is provided for winterized operation. Oil tanks fabricated from 3SO aluminum alloy are mounted on the front face of each firewall. Each tank has a total capacity of 13 gal but the normal flight capacity is 8¼ gal. A slide valve enables oil to reach the engine pump during inverted flight.
While each engine has its separate fuel system, the two are interconnected through a solenoid-operated cross-feed valve so that fuel from any tank is available for either engine in the event of failure. Each system consists of a main tank, a reserve tank and a drop tank and an outer wing leading edge tank. One engine primer supplies both engines.
All equipment used in the fuel system is treated for resistance to aromatic fuel. Main and reserve tanks are located inboard, behind the pilot station, and outboard in the center section between the fuselage and the engine nacelles. The main tank of each system consists of two interconnected units, the outboard main tank with a capacity of 62 gal and the inboard anti-surge tank with a capacity of 31 gal. These tanks are of the self-sealing type. The anti-surge tank contains the fuel gage transmitting unit and the sump. The two connecting fuel lines are fitted with flapper valves at their lower ends to prevent the fuel from flowing back into the outboard main tanks. Each reserve tank is placed between the main beam and the front shear beam. Fuel capacity is 60 gal. The tank is divided by a chord-wise rubber rib to form an anti-surge compartment at the inboard end. A flapper valve in the rib permits fuel to flow into the anti-surge chamber from the main part of the tank but prevents its return.
Early experiments during the evolution of the P-38 produced the streamlined drop tanks which give it an amazing range and which have increased its military utility. The 150 gal drop tanks are suspended from bomb shackles which are hung within fairings from the main beam approximately midway between the fuselage and the booms. The bulkheads of the tanks prevent surging of the fuel from one end to the other. These tanks are formed in two halves from light steel.
Two tank selector valves are located in the lower rear fuselage compartment and manually operated controls, located on the left side of the cockpit, are connected to them by cables. Strainers and rotary fuel pumps are just forward of the valves, and other fuel pumps are located at the aft end of each engine.
A recent addition to the P-38 is the conversion of the outer-wing leading edges to a compartment housing a self-sealing fuel tank of approximately 60 gal capacity in each wing. These tanks are provided with their own electrically driven fuel pumps, strainers, solenoid valves and check valves to connect them into the main fuel system just behind the engine-driven fuel pumps. These tanks are also controlled and selected from the pilot's compartment, thus making it possible to use up the 720 gal of fuel in any desired sequence of tank flow.
Air for the induction system is taken in through scoops mounted on the outboard sides of the forward booms. For desert operation, an air cleaner is mounted in the main wheel wells. By placing it in such a position, an ample sized cleaner could be used without modifying the out- side shape of the plane. By means of a remote control in the cockpit, air coming through the external scoops can be diverted into the wheel well space, from which it passes through the filter and then into the turbosupercharger. Thus, filtered or non-filtered air is available at the discretion of the pilot. Under pressure from the turbosupercharger, it goes to the core-type intercooler underneath the engine, where the heat of supercharging is removed, and is then ducted to the engine carburetor.
The hydraulic system operates the main landing gear and the nose gear, the wing flaps and the coolant radiator exit flaps. Two engine-driven pumps supply and maintain a fluid pressure in the system of 1200-1350 psi. An emergency hand pump with a reserve supply of fluid is provided. The basic system reservoir is installed in the rear fuselage just aft of the radio compartment and strapped to the rear face of the bulkhead. The fluid capacity is 2.1 gal. One Pesco 349GA, self-lubricating hydraulic pump is mounted on each engine accessory drive and is driven at 1.5 times engine speed. These are gear-type pumps with a capacity of 4½ gal per min. A check valve is installed in each pump pressure line so that if one pump fails the other will still be effective, and either pump is capable of maintaining pressure in the entire system. A manually operated disk-type Purolator filter is located between the check valves and the regulator and two integral relief valves are built into the system to prevent stoppage due to fouling of the filter element.
The P-38 installation for actuating the wing flaps is a combination of hydraulic and mechanical power. The actual power is derived from a piston type hydraulic motor which drives the flaps by mechanical linkage. The motor is mounted on the flap drive gear box which is bolted to the center section aft shear beam. A four-way valve, mounted on the right-hand web of the fuselage hydraulic compartment provides control of the direction of flap travel.
The main landing gear and the nose gear have separate but interconnected hydraulic systems and both are controlled by a four-way selector valve mounted on the left-hand web of the hydraulic compartment. The control shaft of the valve extends upward through the floor and is connected by cables and pulleys to the landing gear control lever on the left-hand side of the cockpit.
Lockheed was one of the first manufacturers to turn to the tricycle type of gear as providing optimum conditions for ground contact and handling. With its wide tread, and with the center of gravity well forward of the main alighting wheels, the P-38 handles so well that even inexperienced pilots have no difficulty in controlling it at high speeds on the ground. The level ground position of the plane affords the pilot the best of vision while taxiing, taking off or landing.
All wheels are operated by the main hydraulic system. The gear travel is directly fore and aft about the fulcrum point. When retracted the gear is completely enclosed by flush doors hydraulically operated. Automatic locks are provided for both the up and down positions.
The main landing gear assemblies are located in wells in the front section of the forward booms. The tread is 16' 6" between tire centers. The wheels are magnesium castings equipped with hydraulic brakes and mounting a 36" Goodyear smooth contour tire. The assembly is comprised of an air-oil type shock strut; two drag links, supported by the pivot points; a tubular steel drag strut. one end attaching to the main strut and the opposite end attaching to the drag links; an actuating cylinder, including a "down lock," mounted on the engine mount support bay forward of the firewall and a hydraulically operated "up lock" attached to the inboard side of the forward boom web. The main shock struts are fabricated from X-4130 steel. Arc welded fittings are provided on the cylinder and piston for the attachment of torque arms, drag strut, and the side strut. Two doors, hinged to the lower channels of the forward boom operate automatically with the landing gear movement. A hydraulic cylinder, located in the aft end of the wheel well. operates the front and rear carriage through a linkage of cables and rods.
The nose landing gear consists of a 27" tire. tube and wheel, a fork, two drag struts, cylinder and piston fabricated from X-4130 steel. The gear assembly is kept in alignment, when retracted, by a mechanical centering device incorporated in the internal construction of the piston and cylinder. Alignment is maintained in the extended position by the action of shimmy dampers made up of a fluid reservoir connected with two small spring-loaded hydraulic cylinders mounted on the nose shock strut so that their pistons straddle the shock strut. The supporting bracket encircles and is free to rotate about the shock strut immediately below the drag strut attachment fitting. The aft end of each piston bears on a roller that is carried on a lug integral with the drag strut attachment fitting. Turning the wheel fork forces the piston in on the side toward which the tum is made. The piston's forward stroke is resisted by the hydraulic fluid which must be forced through a small drilled passage in the check valve, while the return stroke is free and rapid, due to the strong extension spring and the automatic opening of the check valve. Through this action an oscillating motion is resisted and progressively reduced.
Noteworthy because of its design, the entire empennage consists of two booms, two vertical stabilizers, two rudders and tabs, one horizontal stabilizer and one tab-equipped elevator.
The elevator is made of one panel and attaches to the horizontal stabilizer by pin-type hinges. The operating cables actuate a torque tube in each empennage boom. These torque tubes are fastened to the elevator screws. The elevator tab is located at the airplane's centerline in the trailing edge of the elevator, and attached by a hinge with a stainless steel hinge pin, which is connected to the actuating unit in the horizontal stabilizer by a push-pull tube. The inboard ends of the torque tubes are carried in bearings on the rear stabilizer spar. The torque tube balance arm, attached to the outboard end of the tube by two taper pins, contains a bearing that slips over a pin to the empennage boom.
Rudders are constructed in two sections and are interchangeable right and left. They are attached to the vertical stabilizer by hinge pins, to the torque tubes and to each other by screws. A counterbalance extends forward of the hinge line of each section. The rudder tab is attached to the rudder by a hinge with a stainless steel hinge-pin, and is connected to the actuating unit by a push-pull tube. The rudder torque tube, bearings, brackets, and arms are assembled as units and are attached to the vertical stabilizer by bolts. The rudder and elevator tab installations are particularly smooth, due to the use of piano type hinges and the flush design of the push-pull control, with a minimum of external protuberance.
The electrical system of the P-38 is so complex that a detailed description is outside the scope of this article. The main wiring diagram lists a total of 143 pieces of electrical equipment. The system is a 24 VDC, single-wire installation with one exception. An inverter supplies 115 VAC for the fluorescent lighting on the instrument panel while the remote compass uses 26 V, 400 cycle AC. Although some wiring is carried in both metal and plastic conduit, much of the system utilizes harnesses which have the advantage of lightness and accessibility.
Besides the ignition systems of the engines themselves, there is the main electrical system which supplies current for the following devices: starters, intercooler actuators, the automatic oil temperature system, propellers, auxiliary fuel pumps, remote compass, panel lights, navigation lights, recognition lights, landing lights, all electric panel instruments, bomb and tank releases, armament firing, cameras, and much other equipment, including the plane's radio transmitter and receivers.
The generator is a 100-amp shunt-wound type driven by the left engine. A voltage regulator in the left-hand main wheel well limits the voltage output to 28.5 V, supplying constant voltage for equipment operation and for battery charging. The main switch box is located directly below the main instrument panel in the cockpit.
The radio compartment is situated immediately aft of the pilot's cockpit and above the main wing beam in the aft end of the fuselage. One of two alternate installations consists of a transmitter and receiver, and a dynamotor unit which is used as transmitter and receiver plate power supply. The remote control box, radio junction box and jack box are on the right hand side of the cockpit within easy reach of the pilot and fuses are located in the junction box.
The other alternate set is a multi-channel aircraft transmitting and receiving unit. It has three ranges of frequency controlled from the cockpit by separate units with ranges of 3 to 6 MHz, 190 - 550 kHz, and 6 - 9.1 MHz, The 190 - 550 kHz band is used as a beacon receiver to receive weather reports, etc. The other two ranges are for military use and are employed for inter-plane communication and similar functions.
Heat is supplied by the engines for both the pilot's cockpit and the armament compartment. Hot air, heated by the right exhaust manifold, is supplied to the cockpit and windshield. An intensifier tube, located within the exhaust manifold just aft of the exhaust "Y" stack in the forward boom, directs the hot air to the butterfly control valve located on the right side of the supercharger. From the butterfly valves, 2" OD tubing passes through the main beam, entering the fuselage aft of the pilot's seat; 1¼" tubing leads forward under the pilot's seat to a slide valve, on the right side of the center control stand. This valve directs hot air on the pilot's feet. A flexible hose carries hot air behind the instrument panel to a tube directing hot air on the windshield. A spot defroster tube stowed in a clip on the left side of the seat may be directed at any part of the canopy or may be clipped to the canopy behind the pilot's shoulder to direct heated air along the top of the canopy. A butterfly valve in the end of the tube controls the flow of hot air.
The left engine supplies the heated air for the armament compartment. The supply system is through a series of tubes similar to those of the cockpit heating arrangement. Within the cockpit a tube passes under the pilot's seat above the fuselage floor to the forward edge of the pilot's seat. From there, it is carried under the floor along the top of the nose wheel well into the left rear of the armament compartment. The outlet is covered with a wire screen. Heat is regulated through controls by a knob on the left windshield frame. As a later modification, electric gun heaters were added, and the output of the left engine intensifier tube was directed into the cockpit.
The oxygen system has been changed from the high pressure system on early P-38s to a low-pressure demand type, three-bottle system. One cylinder is located in the right boom aft of the wheel well and is accessible through a removable panel. The other two cylinders are in the left boom, one just aft of the wheel well and the other one just forward of the battery. All three cylinders are filled from the same filler coupler in the right boom on the forward side of the aft wall of the wheel well. A check valve is located in the filler line immediately preceding each bottle to prevent oxygen from escaping through the filler lines. In the left boom there is a check valve in the supply line near each bottle. In case one bottle develops a leak the check valve will prevent escape of oxygen.
The pressure gage is mounted on the side control stand. A pressure signal is connected electrically to a warning lamp through the airplane electrical system. The signal is adjusted to light the lamp when cylinder pressure drops to approximately 100 psi. The warning lamp and a flow indicator are placed on the central control stand. This installation has an oxygen supply for 6 hr at 30,000 ft. At altitudes less than this the oxygen will last longer due to the automatic mixing feature of the demand regulator. A special mask incorporates an expiratory flapper valve which closes when inhaled permitting oxygen to be drawn into the mask, opening when pilot exhales, permitting exhaled air to discharge into the atmosphere. The regulator mixes pure oxygen and cockpit air in the right proportions for varying altitudes.
The P-38 is heavily armored for pilot protection. Most of the armor plate has been reduced in thickness from the original 3/8" to ¼" since actual combat tests have proved the lighter steel to be fully satisfactory. The armor plate is all face-hardened steel and pieces are kept small to facilitate handling and removal in the field. Each piece is attached separately in its respective place although some overlap to supply added protection. Ten separate pieces make up the installation on the bulkhead aft of the armament compartment which protects the pilot frontally. Two additional pieces protect the cowl just forward of the cockpit canopy and one large plate is mounted directly below the sloping bulletproof windshield. A bent plate completely covers the bottom of the pilot's seat and a flat plate is carried up the back of the seat full width to the lower line of the silhouette armor plate. This plate is mounted at the bulkhead station immediately aft of the pilot and is carried up to the top of the canopy. It is shaped to provide protection for the pilot's head and shoulders but to allow visibility to the rear over his shoulder.
Various alternate gun installations have been tested on the P-38 including groupings of multiple machine guns and cannons never before used on fighter aircraft. Standard installation includes four machine guns and one cannon, although modifications are made outside the factory for special missions and for various theatres of operation.
In the standard grouping there are four .50-cal, type M2 machine guns mounted near the top of the nose. Two are placed on either side of the fuselage center line so that the four guns follow the contour of the top of the fuselage. Type M2 20-mm cannon is mounted on the fuselage center line below the machine guns.
Forward of the fore armament compartment bulkhead a mount is provided for the installation of a type AN N-4 gun camera. The camera is attached to a plate which incorporates a ball and socket joint attached to a bulkhead. Doors in the nose and right hand side of the nose provide access for installation, removal, service and inspection. The camera may be operated either independently or in conjunction with the guns as desired by placing the armament switch in the cockpit on "Camera" or "Combat."
This Design Analysis article was originally published in the August, 1944, issue of Industrial Aviation magazine, vol 1, no 3, pp 7-20, 64-65, 104.
The original article includes a thumbnail portrait of the author, 10 photos, a three-view and 7 detail drawings and diagrams, and 4 data tables, plus a ledger-sized foldout with a color phantom rendering of the plane and diagrams of a number of the planes systems.
Photos are not credited.
Diagrams and drawings: