The L-242 airplane was the result of two concurrent studies which were being pursued prior to Lockheed’s final submission to the OS-130 competition. The first such study was based on the original requirements for a Navy fighter outlined in the first issue of OS-130. The second study was for a day fighter airplane based on requirements for the United States Air Force (which ultimately yielded the legendary F-104). In subsequent changes to OS-130, it became apparent that Navy requirements were paralleling the concept of a day fighter being developed for the Air Force and, as a result, Lockheed’s proposal was essentially the airplane developed for the Air Force modified to fulfill the requirements of a carrier-based airplane for Navy use.
The table is a supplemental summary of the airplane characteristics and performance which describes more clearly the combat capabilities and compares the airplane performance directly with the final requirements of OS-130 and its Amendments.
Lockheed claimed that the L-242 would be entirely adequate to fulfill Navy requirements. It had a top possible combat speed of 1,000 kts, corresponding to a Mach number of 1.74, which was far in excess of the Mach 1.2 required. In addition, it had a combat radius of 495 miles, by the use of droppable fuel tanks and combat at Mach 1.2, which exceeded the 400 miles required.
Since this airplane was identical in general design to the F-104, its internal fuel capacity was not quite sufficient to fulfill the radius specified in OS-130. Thus, the 188 mile radius with a combat speed of Mach 1.2 on internal fuel only was somewhat below the 300 miles required. Lockheed argued that the relatively light weight, coupled with design simplicity and small dimensions, would make the L-242 ideal for carrier-based service.
The previous figures and table show several spot points summarizing the performance of the airplane when the Wright TJ-31-B4 and the General Electric X-24 engine was installed in place of the Wright TJ-31-B3, which was proposed for the basic model. Since these engines could have been installed in the airplane with very little change, and since their development promised to provide a large increase in performance, they provided ample assurance to Lockheed and to the military services that the basic L-242 airframe was not too small to have a large growth potential, both in range and speed, as development engines progressed. With no change except the installation of the X-24 engines, it was apparent that the L-242 could have had a top speed of over Mach 2.0 at 35,000 ft and could potentially have had a combat radius of 550 nm based on a combat number of Mach 1.2 and the use of the same external fuel tanks as then incorporated.
Lockheed emphasized that the L-242 design was a simple airplane based on structural and aerodynamic concepts which had been demonstrated to be valid. Thus, the development of a production airplane would not have been a long process and involved a minimum expenditure for the performance and the utility of the weapon system. The company further emphasized that the development of this model for both the Air Force and Navy would certainly diminish the development costs for both services and provide production costs below those which obtainable for two different airplanes designed for specifications which were so nearly identical.
In the development of the Lockheed L-242 design, the initial effort was spent on an airplane to fulfill the requirements of a non-afterburning fighter which had a combat radius of 300 miles and a high speed of Mach 1. The development of this design was pursued with a substantial effort until November 25, 1952 — in spite of the fact that previous fighter investigations by Lockheed had indicated that this level of performance was quite low and that, on the basis of contemporary power plant data, a lighter smaller, airplane could be achieved using an afterburner type of power plant. These design studies were continued, however, because Lockheed was acutely aware of the Navy’s earnest effort to reduce the size, weight and complexity of its fighter aircraft.
The final design for the non-afterburning fighter did not include a main gear because of the potential saving in weight provided by the mat landing system.
It was gratifying, in the comparison of the airplanes, to see that the supersonic design was not vastly larger, and was only more complicated by the inclusion of the afterburner. It was the considered opinion of Lockheed that the final requirements were more realistic and have resulted in a more useful airplane for the Navy.
The following sections summarize the operational characteristics of the airplane, including details of combat performance on alternate missions, combat armament, carrier suitability and adaptability to later designs of carrier landing facilities. In addition, its maintenance features and emergency escape provisions are outlined. Another section is devoted to the development background for this particular design, with discussions on why the various design details were created. Finally, detail descriptions of each of the airplane basic components are given along with a section on the design for producibility and a summary of the weight.
General Airplane Description
The general arrangement and inboard profile provide the complete description of the airplane from a configuration standpoint. Of particular interest was the fact that a simple, thin, straight wing was employed, mounted in a mid-position on a relatively symmetrical fuselage with side ducts. The cockpit was mounted well forward in the fuselage for visibility and to keep the volume of the canopy from being concurrent with the volume of the ducts on a lengthwise distribution of fuselage volume thereby obtaining low fuselage drag in the transonic and supersonic regions.
The vertical and horizontal tail surfaces were developed in the wind tunnel and represented very nearly the optimum in terms of area and position to obtain linear stability characteristics and adequate high and low speed control. The high position of the horizontal tail was found to be much superior to any other position.
In the inboard profile, it is worth noting that the ammunition and fuel were placed in a position as close to the airplane CG as possible to ensure that good stability and control were achievable at any loading condition. It should be pointed out that no fuel was placed in the wing although the simple structure and heavy skin made this an ideal stowage space for an additional 50 gallons. The wing tanks were avoided because they increased the vulnerable area of the airplane by orders of magnitude so that the utility of the weapon system was diminished rather than increased by the addition of this internal fuel.
Also apparent on the inboard profile was the fuselage stowage of the main gear, which permitted the 3.4% thickness of the wing. In this location, the gear did not interfere with the external wing stores nor did it prohibit the installation of external stores on the fuselage.
Finally, as shown in both Figures A and B, the ducts and the speed brakes had both been placed on the side of the fuselage so that eventual modification of the airplane for mat landing purposes would not involve any change in the aerodynamic configuration.
The general arrangement emphasized the simplicity of the airplane concept which was necessary to obtain a really useable military weapon and to decrease the development, manufacturing and operational costs of the weapon system.
Since the definition of a combat profile was difficult and the definition of the combat portion of a mission could be interpreted in many ways, Lockheed created a summary of alternate combat profiles to illustrate the full utility of the L-242. Several alternate combat profiles are shown which were based on different missions where equipment was changed, or the mission purpose was changed, from that of air-to-air fighter. Included in these alternate missions were photo missions, bombardment operations, and trainer uses, all of which were found to be easily provided by simple adaptation of the basic design.
Since the operation of a supersonic airplane had lead to many conjectures concerning the limits of the airplane and its utility in combat maneuvers, considerable emphasis was place on the ability of the airplane to accelerate rapidly to combat speeds, and on its ability to decelerate from these speeds in order to obtain speed flexibility. Although the combat Mach number requested by the Navy was only 1.2, the inherent high-speed capability of the airplane of Mach 1.74 provided extremely rapid acceleration to obtain any desired Mach number in a more modest range. The Navy combat Mach number could be achieved from level flight cruise in 65 seconds. Furthermore, a deceleration from the Navy Mach number of 1.2 to a Mach number of .90, which was the best climb speed of the airplane, could be effected in 20 seconds.
Since the Navy requirements called for a combat radius of 300 miles with full internal fuel, and since it was recognized that the Navy had an extreme logistics problem with droppable fuel tanks, it was felt to be desirable to investigate alternate tip tank installations which would be of the permanent type to determine what the performance of the airplane would be with these permanently installed tanks.
A combat radius of 300 nm could be achieved with full internal fuel by the addition of permanent tip tanks containing 60 gallons each, for which a penalty of .05 Mach number was paid in high speed performance. It was recommended that permanent tip tanks not be installed in the prototype phases of this program since their size would depend on mission definition and since acceleration to high speed suffered substantially. It should be emphasized that a mat landing version of the airplane did not have sufficient internal fuel to achieve the desired radius of 300 miles.
The basic Mark 16 sight installation, coupled with AN/APG-34 range radar, was believed to be the simplest effective installation which could have been incorporated in the first models of the L-242. Since the components of this sight and system were essentially developed, and since training procedures were in existence which would develop pilots for this system, it was felt that its installation in first-production models of the airplane was certainly justified. With all of the armament components which were proposed for this system, with the exception of the Aerowolf missile, it was believed to be the simplest possible sight arrangement.
The accompanying blueprints summarize the armament investigations which were made on the airplane and showed the following capabilities.
- Gun Installations—As shown in Figure 9, the basic Mark 12 guns were feasible to install in the airplane with 500 rounds of ammunition as required. Alternatively, two T-182 cannon could have been installed with 500 rounds of ammunition, or a single T-171 cannon could have also been included with 500 rounds of ammunition. Lockheed believed that the use of the T-171 cannon was worthy of substantial consideration, since it had the capability of firing almost as many rounds per second as the four Mark 12 guns, and its installation utilized only one armament compartment on the airplane, leaving the other armament compartment available for the installation of additional equipment or rockets. A comparison of the total armament weight for the several guns indicated that for a given amount of fire power there was little to choose between the installations from a weight standpoint.
- Rocket Installations—Lockheed only presented one type of rocket installation, although other types could have been installed on the airplane, and the installation of this rocket is shown in Figures 10 and 11. On Figure 10, the installation of rockets in the fuselage was shown to consist of two openable rocket boxes which were installed in placed of the guns. These were supplemented by an installation of rockets in the nose of the airplane which would have required some rearrangement of the equipment. Since the basic Navy specification stated firmly that rocket installations had to be made with open tubes, it was worthy of note that these installations deviated from the requirement for what were believed to be substantially good reasons. The rocket round launched from a closed tube had been determined to have almost 100% increase in launching velocity. Since the accuracy of the rocket was affected by damping provided by the fins which diminished with altitude and improved with velocity, and since accuracy was also affected by the airflow around the airplane body and the load factors being imposed on the airplane at the time of launching, it was extremely desirable that the rocket attain the highest possible velocity and get away from the airplane influence in the soonest possible time for maximum accuracy. The closed tube was apparently a very promising way of obtaining these results. Tests at Lockheed comparing closed and open tubes were then in process, and there was no evidence to show that a closed tube had any detrimental effect on the flight of the rocket. As shown in Figures 10 and 11, both closed and open tubes were used, depending on the installation desired. In the fuselage, it was obvious that the closed tube mechanism was desirable, aside from advantages obtained in rocket accuracy and, therefore, only the closed tubes were shown. For the pod mounted rockets which could have been added to the airplane without changing the basic armament, it was immaterial whether closed or open tubes were used and, in fact, it could have been even desirable to have open tubes on the tip pods, accepting loss in range and accuracy for the ability to put rockets in tandem. In any case, the sight had to be arranged for either open or closed tubes, and, therefore, it would have been undesirable to combine the fuselage rocket with tip rocket pods, since the ballistics of the two would have been different. Lockheed believed that the most desirable installation could have been a combination of fuselage guns and tip rockets; or a combination of fuselage guns and nose rockets. A final alternate would have been the use of a T-145 gun on one side with rockets on the opposite side in the armament compartment. This was the lightest combination of rockets and guns, and could have proven the simplest.
- Missile Installations—Figure 12 shows the installation of four Sparrow missiles and four Aerowolf missiles on the airplane mounted on external racks on the wing. The mid-wing arrangement made these installations relatively simple with no ground clearance problem and the straight wing permitted their installation with the least effect on balance. Lockheed gave special consideration to the airplane with the Aerowolf, since this missile promised the simplest possible sighting arrangement, eliminating range determination by the sight and eliminating the necessity of lead computation. Lockheed believed usage of the Aerowolf missile as sole armament merited special consideration for versions of the airplane which were to do interceptor duty only. This would have permitted additional fuel in the fuselage, resulting in increased radius or combat time.
- External Bomb Stores—Since the development of a fighter store was far from firm, Figure 13 was compiled to show several alternative arrangements of external stores on the airplane. It was worthy of note that the mid-wing arrangement of the airplane and the landing gear installation had placed substantial emphasis on the ability of the airplane to carry external stores with good ground and landing gear clearance. As shown on Figure 13, the 1,000 lb Douglas store was easily carried, and a store of comparable shape with a diameter of 14″ and folding fins could have been carried at the centerline of the airplane, thus permitting a high-weight store without the necessity of eccentric lateral weight.
- Alternate Fire Control System—An alternate radar fire control system which replaced the AN/APG-34 with a new type radar had been studied and appeared to have substantial merit. This device would have eliminated the main operational limitations imposed by installing the AN/APG-34 ranging radar with the Mark 16 optical computing sight in the Lockheed Model L-242 fighter. Furthermore, it appeared to be feasible with negligible increase in equipment weight and complexity. This system included a simple lightweight search radar combined with the Mark 16 sight. To justify a change, an examination of the inherent deficiencies in the APG-34 radar was in order. Its primary disadvantage was that it was designed to supply radar range information “only” to an optical computing sight and therefore had essentially no target detection capability. In contemporary jet fighter aircraft, flying at supersonic speeds, it was believed desirable to extend detection of fighter targets to ranges of 10, 12 and possibly 15 miles. Since detection of these same targets by visual means was definitely limited, it followed that a search radar with these capabilities could have been employed if it were simple enough and had a simple presentation. The proposed radar was basically a simple single-purpose 100 KW “X” band pulse type search set with one sweep range, 25 nm. Detection information including range and approximate target direction were simultaneously displayed to the pilot on a 5″ diameter high-intensity cathode ray tube indicator. With this information the pilot was able to fly his aircraft and track the target to a visual detection range. This was accomplished without target lock-on and automatic tracking circuitry was was used in contemporary systems such as the AN/APQ-41 and the AN/APG-37. Once visual contact was made the radar was used to supply target range information to the Mark 16 optical computing sight from which gunfire was directed. A receiver-transmitter modulator, antenna, indicator synchronizer, and control box were included in the proposed radar. The following is a brief itemized operational description of the system: a) Controls—Operation was instigated by the pilot who had one primary control, an ON-OFF switch. Focus, brilliance and receiver gain controls were added so that the picture could be adjusted for optimum viewing. These latter controls could be set on the ground during pre-flight checks. b) Antenna Scan Coverage—The search antenna, a paraboloid approximately 22″ in diameter, mounted in the aircraft’s nose section, revolved at a high rate of scan in a circular pattern about the airplane’s centerline covering a conical area of plus/minus 60° from this same center line. This coverage was accomplished by mechanical movement of the antenna as it spun in either a spiral or nodding manner. c) Indicator Presentation—The cathode ray tube indicator picture which was used by the pilot for search and tracking purposes was a vertical type PPI display which revolved in synch with the antenna’s rotation. In other words, the fighter airplane was at the center of the scope and the sweep gave range and approximate elevation and azimuth position information of the target. This is further illustrated in Figure 14. Roman numeral I of Figure 14 shows how this presentation would have appeared, based on a sweep range of 25 nm. The earth’s radar return would have essentially formed an artificial horizon at the bottom of the scope and further given the pilot his approximate altitude and attitude above the earth’s surface. Reference marks showing range, as well as horizontal and vertical reference lines were etched on the face of the cathode ray tube. Targets that appeared above the horizontal line were those targets above the fighter and those which appeared to the right of the vertical etched line were to the right of the airplane and those that appeared on the left of this line were to its left. Thus, if a target was detected as shown in Roman numeral II, it was up and to the right at a range of 15 miles. The pilot’s next move was to fly in the general direction of this target as shown in illustration III. When the target was approximately on the airplane’s center line it broke into a circle as shown in illustration IV. The circle resulted from the fact that the antenna beam was illuminating the target in all of its axial rotational positions. The pilot then flew the presentation keeping a circle on the scope at all times. This meant that he was tracking the target accurately. As the range reduced, the circle shrank in size until he could visually see the target. At this point, the radar could be locked on to the target in range only and would supply this range information to the Mark 16 sight. From this point to the kill, tracking and subsequent gunfire was identical to the normal operation with only a range radar. It was believed to be highly desirable to develop such a fire control system to improve the utility and target acquisition range of the day fighter airplane. If the principles of operation outlined above were strictly adhered to and no additional requirements were imposed on the radar system, Lockheed believed that a truly simple system could have been devised.
- Minimum Production Man-Hours—Man-hour savings had been accomplished by keeping the number of parts in the airplane to a minimum. For example, the fuselage structure used no stringers; just skin and rings and three main longerons. Skins for the box section of the wing were one-piece units, thus eliminating joints and their related attachments. The primary wing beams attaching to the fuselage were single piece forgings at the root. The fuselage sections broke down into half-shell side panels affording unlimited workman access and permitted pre-installation of functional and equipment items to the fullest extent. Accommodations in the structure design, for cable, wiring, and piping runs were made to provide a maximum of access for initial installation as well as for service. The electronic equipment was designed as a package unit to facilitate installation, maintenance, and service.
- Economy of Tooling and Floor Space—Simplified tension type joints used throughout the fuselage, and for joining the wings, kept master gauge costs to a minimum and eliminated the need for any mating jigs. Each major section of the airplane was designed to be assembled complete as a subassembly prior to mating. For example, the forward fuselage, Station 122 to 378.5, could have been assembled complete, including the canopy, nose landing gear, all doors and panels, etc. prior to mating to the mid-fuselage. All doors, panels, etc. fell within the confines of each major subassembly. This feature eliminated the need for portable apply-type drill fixtures and any drilling or fitting operations after mating that was normally required for members straddling joint planes. The final assembly line for the L-242 would have been comparatively short since the operation would have consisted primarily of mating the major sections, with installation work kept to a minimum.
- Sub-Contracting the Expensibility—Each component section of the L-242 was a complete unit within itself, and was particularly adaptable for subcontracting. All units were small enough to handle and ship with ease. All problems of tool coordination were confined to a single section since all movable or removable members were contained in the subsection. Single plane tension joints, and their related simplified tooling, assured proper fits and mating of the various sections of the ship. The units of the production breakdown were extended sufficiently to permit proper work distribution for expanded production programs. Only the duplication of tools and facilities required application to attain efficient mobilization rates. The manufacturing methods and techniques used in the production of the L-242 were standard methods of manufacture and were all used on contemporary Lockheed production aircraft.
- Forward Fuselage Stations 122 to 378.5—The forward fuselage was composed of three major components; the nose section and the left and right side panels. Each unit offered complete freedom for pre-installation. Mating was accomplished by splicing the four bulkheads at the centerline, attaching the bottom center longeron, joining the nose section at the front pressure bulkhead, and framing in the opening for the pilot’s escape hatch. The forward fuselage was joined to the mid-fuselage by tension bolts at the three longerons and a series of tension screws around the periphery. The nose radome was attached by three tension bolts. The electronic equipment, which was located directly aft of the cockpit, as shown in Figure 26. This rack could be raised for unlimited access in service and check out. For initial assembly, the entire unit would have been built up complete, aligned, and checked out prior to being installed in the airplane. The canopy, the air scoops, the nose landing gear, and all doors and panels were completely installed in this subsection prior to mating with the midsection.
- Mid-Fuselage Stations 378.5 to 508—The mid-fuselage consisted of three main sub-assemblies; the upper panel, and the left and right side panels. Accompanying these units were the bottom center longeron, the firewall bulkhead, and the rings at Station 378.5 to 508. These fully accessible subassemblies were mated by joining the panels at the three longerons, and installing the firewall bulkhead and the two end rings. The main landing gear, the main landing gear doors, the tank panel, and engine access doors were installed complete in this section prior to mating to the forward fuselage. The fuel cells and the engine air ducts were installed after mating, to provide access to the joint. The mid-fuselage to aft fuselage joint was a quick disconnect three bolt tension joint similar to those features on previous Lockheed jet fighter planes.
- Aft Fuselage Stations 508 to 618—The aft fuselage structure broke down into six skin and ring segment panels, three main longerons, and two empennage attaching bulkheads. This unit was assembled by framing the longerons to the bulkheads and the joint ring at Station 508, and attaching the panel assemblies at the longerons. The tail cone joint, at Station 618, was a simple tension type joint accessible from inside the fuselage. This section, along with the empennage, tail cone, and dive flaps, constituted the service removable aft section, for engine access and exchange.
- Wing—The wing, which featured single unit skins and a series of spanwise beams, attached to the side of the fuselage with a simple flat plane tension joint. Twenty-one attaching bolts were accessible from inside the fuselage. The breakdown of the wing is shown in Figure 27. All movable surfaces were to have been installed on the wing, along with their respective actuator units, and rigged and checked out prior to being mated to the fuselage. All control mechanisms for actuating the leading edge and trailing edge flaps were located outside the primary wing box structure, thus eliminating removable access panels in the box structure. Access was gained by disconnecting the flap links and allowing the surface to hinge downward approximately 90°. The well, which housed the spoiler in the wing surface, was designed to be sub-assembled complete with spoiler, cylinders, piping, etc., prior to being installed in the wing structure.
- Empennage—The empennage was attached to the aft fuselage by four shear bolts at the vertical fin beams. Access to this joint was gained through removable fillet sections. The empennage contained the pitch control booster mechanism and was capable of being completely rigged prior to mating to the fuselage. a) Vertical Fin—The vertical fin was composed of a leading edge assembly, a trailing edge assembly, a tip, and intermediate ribs and skins. The two spars, which were aluminum forgings, had integral fuselage attaching lugs and the rear spar could have had an integral trunnion bearing for the stabilizer. This unit was assembled by positioning the leading edge and trailing edge, and framing in the intermediate ribs and skins. The rudder hinges where designed to take all end loads at one hinge thus providing clearance at the other two hinges. The fin tip and the upper trailing edge section were removable, to provide access to the stabilizer controls and the rudder hinges. b) Horizontal Stabilizer—The stabilizer consisted of a leading edge assembly, a trailing edge assembly, a center, and tip rip, tip fairings, a dual main beam, and intermediate skins and spars. This structure was very simple since there were no elevator complexities to deal with. The center ribs were simple aluminum forgings. The trailing edge section featured a metal bonded honeycomb core for stiffening. The spars were simple extruded or sheet metal sections. The single unit design of the ribs and spars kept assembly operations to a minimum and involved only simple and straight forward fabrication operations.
The design was based on a background of flight experience and wind tunnel experience which covered transonic and supersonic aircraft with straight, swept and delta wings so that the resulting configuration could have been considered an optimum solution based on known facts.The straight wing L-242 was believed to have had the best attainable low-speed flight characteristics for an airplane whose high speed flight potential approached Mach 2. The basic airplane design provided for normal operational procedures, normal maintenance procedures and utilized normal components of armament, electronic and mechanical aids to fill all of the system performance necessities. This assured the Navy that a fully operational fighter system would be available for a minimum of cost in the shortest possible time. In spite of the conventionality of the airplane design, no performance penalties had been incurred, and it was entirely feasible to modify the airplane for use with an unconventional mat landing system with attendant gains in combat time or radius.
Although the basic design had been limited to a fighter task, and even though a concentrated effort had been spent reducing the size, complexity and weight of the airplane, Lockheed believed the design was usable for several alternate missions including photo reconnaissance, training and bombardment missions, even including the carrying of special stores. In spite of the high speed potential, pilot emergency escape was entirely feasible with simple foolproof mechanisms, thus aiding substantially in maintaining pilot morale. Combining a high speed potential of Mach 1.74 (1,000 kts) with a combat altitude of 50,000 ft and combat radii varying from 188 miles to 495 miles, depending upon fuel carried and type of mission, the L-242 had the potential of performance to give air superiority or equality over any land-based fighter that could have been developed in the equivalent time period. Furthermore, it had the potential with contemporary programmed power plants of even higher performance and greater range and these power plants could have been installed with very minor changes to the basic airframe. Since the basic L-242 airplane was under development for use on a similar mission by the Air Force, its concurrent development for Navy purposes was a logical program. It provided both services with a maximum striking force for minimum development and production costs. Since the requirements as outlined by The Air Force and by the Navy were roughly equivalent, it appeared uneconomical to develop two airplanes for such similar missions.
Figure 3 at the beginning of this article covered the performance of the airplane in and around the carrier, but did not emphasize the mechanical suitability and the arrangement of the launching and arresting gear on the airplane. Figure 15 is a diagrammatic outline of the hook and launching arrangement showing that all the Navy requirements concerning these features had been met. Note that this drawing appears to show the early, non-afterburning version of the design.
With regard to carrier operation, Lockheed believed that the L-242 was completely conventional and would have had no peculiarities requiring special operational techniques. Of particular interest was the fact that the nose of the airplane had been kept as narrow as possible in an effort to improve the visibility from the cockpit. Since it was difficult, if not impossible, to obtain ideal vision directly through the canopy forward at high angles of attack, Lockheed felt that this portion of the canopy should be made as narrow as possible and the nose be correspondingly narrow in an effort to provide vision to both sides, where the angle of incidence between the pilot’s vision line and the glass was as large as feasible. The windshield design achieved this characteristic by making the forward portion of the windshield of cylindrical shape, cut only by the very narrow flat required by the sight.
Figure 16 is a diagram of the L-242 spotted on the deck space outlined in OS-130 with the largest tip tanks installed. Twenty-five airplanes were possible to spot in this area without resorting to any unconventional or peculiar maneuvering to permit all airplanes access to the take-off area.
One particular item of OS-130 suggested that the design of the airplane be adaptable to the use of mat landing. For this reason, special emphasis was placed on the design of the airplane to keep the wing tips and wing high, to keep the dive brakes off the belly of the airplane and to put the landing gear in such a position that the space made available by the removal of the gear could have been utilized for alternate purposes.
Figure 17 is a diagram of the airplane inboard profile in which the landing gear has been removed providing additional space and saving weight. It appeared that approximately 100 gals of fuel could have been put in this area, thereby extending the radius to the desired 300 miles. Lockheed suggested that when such a version of the airplane was considered the nose gear be retained so that the cart for maneuvering the airplane on the carrier decks would only have to be of the two-wheel variety once the nose gear had been extended. This could have been of substantial assistance for take-off purposes and may have simplified the carrier procedure. Furthermore, it would have permitted an external gear to be mounted on the wing, thus facilitating the use of the airplane for training and ferrying purposes where the mat landing facilities were not available.
Figure 18 illustrates the provisions which were made for emergency egress from the airplane. Considerable effort was spent analyzing the various means of exit from a high-speed airplane including capsule exit, nose detachment, upward and downward ejection, including the actual detail design of several of these alternates. After substantial consideration of this problem, it was concluded that the simplest and, therefore, the most feasible method of high-speed exit, resulted from a combination of complete airplane deceleration and downward ejection.
Experiments at the Lockheed Corporation under the development of the MX-883 supersonic ram jet test vehicle had shown that the deceleration by parachute of large-weight vehicles was entirely feasible and practical, and therefore, should have been seriously considered because of it simplicity. As shown on Figure 18, it was proposed that either a 11 ft diameter drag chute or a reefed 16 ft diameter drag chute be incorporated on the airplane for this purpose. The compartment had to be large enough to contain a 16 ft chute, since it was useful also as a landing deceleration device when the airplane was used on normal airports. For carrier use only, it could have been desirable to include only a drag chute, since it was smaller and could have been more dependable. With either parachute, the airplane could have been decelerated from a Mach number of 2 in level flight at combat weight to a safe speed for ejection in 7 seconds with a loss of less than 1,000 ft of altitude.
The downward ejection simplified the design of the cockpit, since knee clearance to the instrument panel and canopy no longer were necessary. Furthermore, the opening and closing mechanisms of the canopy could have been simplified, since it did not have to open at high speeds. For ditching purposes, of course, the canopy was easily opened and removed. Lockheed believed that the shape of the L-242 and its smooth belly configuration with a high wing would have permitted excellent and smooth ditching with a minimum of damage to the airplane.
Figure 19 is a breakdown of the airplane and its manufacturing components showing that the basic design followed conventional practice which had been developed over very long periods of time in designing useful and easily maintained airplanes. Of particular interest was the fact that each component was relatively small and easily handled. This was particularly true of the wing panels which could be detached from the sides of the fuselage and replaced in a simple fashion. A further advantage of the configuration was that no wing folding mechanism was necessary because of the short span, thereby eliminating one major item from maintenance consideration.
Figure 20 is a diagram showing the access doors which were provided for servicing the various components of equipment installed in the airplane. Unusual in this regard was the fact that the electronic equipment was concentrated in two easily accessible locations with large doors, and the wheel well doors within the fuselage provide additional access to portions of the fuel system and hydraulic system of the airplane. Removing the tail for engine access was a completely accepted maintenance feature with many advantages as pioneered in the F-80 series airplanes. A three-point attachment of the aft fuselage simplified its removal, and rails were provided in the engine installation to permit the movement of the engine aft out of the fuselage to such a position that the hoist attachment could have been made directly without any special rigging. The engine was removed as an entity, including its afterburner. Of further interest to maintenance was the fact that single-point refueling was incorporated in the design which permitted rapid and simple refueling procedures, cutting down turn-around time.
Development Background for Design
Since a number of reasons for the airplane configuration were already outlined in the discussion of the operational characteristics of the design, it is only necessary to make a brief recapitulation of these reasons. There was a wealth of development background which included the following major items of interest.
The development of an optimum wing for supersonic flight had been the subject of a vast amount of research and design investigation at Lockheed, and substantial confirmation existed for the selection of the straight wing on the L-242.