Lockheed L-2000

The Lockheed L-2000 was Lockheed's entry in a government-funded competition to build the United States' first supersonic transport (SST) in the 1960s. The L-2000 lost the contract to the Boeing 2707, but that competing design was ultimately canceled for political, environmental and economic reasons.

In 1961, President John F. Kennedy committed the government to subsidizing 75% of the development of a commercial airliner to compete with Anglo-French Concorde then under development. The director of the Federal Aviation Administration (FAA), Najeeb Halaby, elected to improve upon the Concorde's design rather than compete head-to-head with it. The SST, which represented a significant advance over the Concorde, was intended to carry 250 passengers (a large number at the time), fly at Mach 2.7-3.0, and have a range of 4,000 nautical miles (7400 km).

The program was launched on June 5, 1963, and the FAA estimated that by 1990 there would be a market for 500 SSTs. Boeing, Lockheed, and North American officially responded. North American's design was soon rejected, but the Boeing and Lockheed designs were selected for further study.

Early design studies

Like Boeing, Lockheed had done a number of "paper studies" on various SST designs, starting in 1958. Lockheed sought an airplane with cruise speeds of around 2000 mph with takeoff and landing speeds that compared to large subsonic jets of the same era. They also desired a plane whose center of lift (C/L) could be managed throughout the entire speed-range. Lockheed knew a variable-geometry swing-wing could accomplish this goal, but felt it was too heavy; they wanted a fixed-wing design that could accomplish this goal. In a worst case scenario they were willing to design a fixed-wing aircraft using fuel for ballast.

Early designs followed Lockheed's tapered straight-wing much like the type used on the F-104 Starfighter, with a delta-shaped canard for aerodynamic trim. The problem was that in wind-tunnel tests the shift in the airplane's C/L was substantial, and a delta wing was substituted which helped some, but was not deemed sufficient. By 1962, Lockheed arrived on a highly-swept bat-wing design featuring four-engine pods buried in the wings, and a canard. The improvement was closer to their goal, but not optimal. By 1963 they extended the leading edge of the wing forward a bit to eliminate the need for the canard, and re-shaped the wing into a double-delta shaped with a mild twist and camber. This, along with careful shaping of the fuselage, was able to control the shift in the center of pressure caused by the highly-swept, forward part of the wing developing lift supersonically. The engines were shifted from being buried in the wings to individual pods slung below the wings.

The airplane was designated L-2000-1 and was 223 feet (70 m) long with a narrow-body 132 inch (335.2 cm) wide fuselage to meet aerodynamic requirements, allowing for 5-abreast coach-seating in coach and 4-abreast in first-class seating. A typical mixed-class seating layout would equal around 170 passengers, with high-density layouts exceeding 200 passengers.

The L-2000-1 featured a long pointed nose which was almost flat on top and curved on the bottom which allowed improved supersonic-performance, and could be drooped for takeoff and landing to provide adequate visibility. The wing design featured a sharp forward inboard sweep of 80 degrees, with the remaining part of the wing's leading edge swept back 60 degrees, and an overall area of 8,370 square feet (778 sq m). The high-sweep angles produced powerful vortices on the leading edge which increased lift at moderate to high angle of attack, and still retained stable flow over the control surfaces during a stall. These vortices also provided good directional control as well, which was somewhat deficient with the nose-drooped at low-speeds. The wing, while only 3-percent thick, provided substantial lift due to its large area, aided by vortex lift, and allowed takeoff and landing speeds which compared to a Boeing 707. Additionally, a delta-wing is a naturally rigid structure which requires little stiffening.

The plane's undercarriage was a traditional tricycle type with a twin-wheeled nose-gear, and with each of the two six-wheeled main legs utilizing the same tires used on the Douglas DC-8, but filled with nitrogen and at lower pressures.

To provide an optimum entry date into service, Lockheed decided to use a beefed-up turbofan derivative of the Pratt & Whitney J58. The J58 had already successfully proven itself as a high-thrust, high-performance jet engine used on the Top Secret Lockheed A-12 (and subsequently, on the Lockheed SR-71 Blackbird.) Due to it being a turbofan, it was deemed to be quieter than a typical turbojet at low altitude and low speed, requiring no afterburner for takeoff, and allowed reduced reduced power settings. The engines were placed in cylindrical pods with a wedge-shaped splitter, and a squarish intake providing the inlet system for the aircraft. The inlet design was designed with the goal of requiring no moving parts and was naturally stable. To reduce the noise from sonic booms, rather than penetrate the sound barrier at a more ideal ~30,000 feet, they intended to penetrate it at 42,000 feet instead. It would not be possible on hot days, but on normal days this would be achievable. Acceleration would continue through the sound barrier to Mach 1.15, at which point sonic booms would be audible on the ground. The plane would climb precisely to minimize sonic boom levels. After an initial level-off at around 71,500 feet (21.8 km), the plane would cruise-climb upwards, ultimately reaching 76,500 feet (23.3 km). Descents would be also performed in a precise way to reduce sonic boom levels until subsonic.

By 1964 the US Government issued new requirements regarding the SST Program which required Lockheed to modify their design, by now called the L-2000-2. The new design had numerous modifications to the wing; one change was rounding the front of the forward delta in order to eliminate the pitch-up tendency. To increase high-speed aerodynamic efficiency, the wing's thickness was reduced to 2.3 percent, the leading edges were made sharper, the sweep angles were changed from 80/60 to 85/62, and substantial twist and camber were added to the forward delta, while much of the rear delta was twisted upwards to allow the elevons to remain flush at Mach 3.0. In addition, wing/body fairings were added on the underside of the fuselage where the wings are located, allowing a more normally-shaped nose to be used. To retain low-speed performance, the rear delta was enlarged considerably, and to increase the payload the trailing edge featured a forward sweep of 10 degrees. The new nose reduced the overall length to 214 feet (65.2 m) while retaining virtually the same internal dimensions. Wingspan was identical as before, and despite the thinner wing, the increased wing-area of 9026 sq. feet (838.5 sq m) allowed the same takeoff performance. The airplane's overall lift-to-drag ratio increased from 7.25 to 7.94.

During the course of the L-2000-2 development the engine previously selected by Lockheed was no longer deemed acceptable. During the time-frame between the L-2000-1 and L-2000-2, Pratt and Whitney designed a new afterburning turbofan called the JTF-17A which produced greater amounts of thrust. General Electric developed the GE-4 which was an afterburning turbojet with variable guide-vanes, which was actually the less powerful of the two at sea level, but produced more power at high-altitudes, both engines requiring some degree of afterburner during cruise. Lockheed's design favored the JTF-17A over the GE-4, but there was the risk that GE would win the engine competition and Lockheed would win the SST Contract, so they developed new engine pods that could accommodate either engine. Aerodynamic modifications allowed a shorter engine pod to be used, utilizing an inlet with minimal external cowl angles and precisely contoured to allow a high-pressure recovery using no moving parts and allowed maximum performance with either engine option. To allow additional airflow for noise-reduction, or to aid afterburner performance, a set of suck-in doors was added to the rear portion of the pod. To provide mid-air braking capability for rapid deceleration and/or rapid descents, and to assist ground braking, part of the nozzle can be used as a thrust reverser for use at speeds below Mach 1.2. The pods were additionally repositioned on the new wing to better shield them from abrupt changes in airflow.

The additional thrust from the new engines allowed supersonic penetration to be delayed until up to 45,000 feet (13.7 km) under virtually all conditions. Since at this point the possibility of supersonic overland flight was considered to be an option, Lockheed considered larger, shorter-ranged versions of the L-2000-2B's. All designs weighed exactly the same, with a new tail design, changes to the fuselage length, extensions to the forward-delta, increased capacity, and variations in fuel capacity. The largest version featured capacity for 250 domestic passengers, while the medium version featured transatlantic capability with 220 passengers. Despite the fuselage length changes, there was no appreciable increase in the risk of the aircraft pitching upwards too far (over-rotation) on takeoff.

Design competition

By 1966, the design took on its final form as the L-2000-7A and L-2000-7B. The L-2000-7A featured a re-designed wing and fuselage lengthened to 273 feet (83 m). The longer fuselage allows for a mixed-class seating of 230 passengers. The new wing featured a proportionately larger forward delta, with greater refinement to the wing's twist and curvature. Despite the same wingspan, the wing-area was increased to 9424 square feet (875 sq m), with a slightly reduced 84 degree sweepback, and an increased 65 degree main delta wing, with the trailing edge's forward sweep reduced. Unlike previous versions, this aircraft featured a leading edge flap to increase lift at low speeds, and to allow a slight down-elevon deflection. The fuselage, as a result of greater length, changes to the wings design, and attempts to reduce drag further, featured a slight vertical thinning in the fuselage where the wings are, a more prominent wing/body "belly" to carry fuel and cargo, longer nose, and a refined tail. Since the airplane was not as directionally stable as before, the plane featured a ventral fin, located on the underside of the trailing fuselage. The L-2000-7B was extended to 293 feet (89 m), utilizing a lengthened cabin and a more pronounced upward-curving tail to reduce the chance of the tail striking the runway during over-rotation. Both designs had the same maximum weight of 590,000 lbs (267,600 kg), and aerodynamic lift-to-drag ratio increased to 8:1.

Full-scale mock-ups of the Boeing 2707-200 and L-2000-7 designs were presented to the FAA, and on December 31, 1966 the Boeing design was selected. The Lockheed design was judged simpler to produce and less risky, but its performance during takeoff and at high speed was slightly lower. Because of the JTF-17A, the L-2000-7 was also predicted to be louder as well. The Boeing design was considered more advanced, representing a greater lead over the Concorde and thus more fitting to the original design mandate. Ironically, Boeing eventually changed its advanced variable-geometry wing design to a more simple delta-wing similar to Lockheed's design, but with a tail. If Lockheed had built its simpler design, it might have flown by 1971. With technical problems, delays, cost overruns, and environmental and economic questions, the Boeing SST was ultimately canceled on May 20, 1971 after the US Congress stopped federal funding for the SST program on March 24, 1971.

See also

• List of Lockheed aircraft

References

• Boyne, Walter J, Beyond the Horizons: The Lockheed Story. St. Martin's Press: New York, 1998.
• Francillon, René J, Lockheed Aircraft since 1913. Naval Institute Press: Annapolis, 1987.