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Boeing 2707 SST Design

Although supersonic flight was first accomplished in 1947 by Captain Charles E. Yeager flying the X-l airplane, serious study of the problems of long-range, supersonic cruising aircraft did not begin until the mid-1950's. Research on such aircraft had its origins in the United States in the work which was begun in 1954 or 1955 in support of the Air Force XB-70 program, and research aimed toward a supersonic commercial airliner began in 1958.

A bow shock wave exists for free-stream Mach numbers above 1.0. In three dimensions, the bow shock is in reality a cone in shape (a Mach cone) as it extends back from the nose of the airplane. The Mach cone becomes increasingly swept back with increasing Mach numbers. As long as the wing is swept back behind the Mach cone, there is subsonic flow over most of the wing and relatively low drag. A delta wing has the advantage of a large sweep angle but also greater wing area than a simple swept wing to compensate for the loss of lift usually experienced in sweepback. But, at still higher supersonic Mach numbers, the Mach cone may approach the leading edge of even a highly swept delta wing. Because of the high sweep, the landing speeds of airplanes with delta wings are very fast.

Sweepback has been used primarily in the interest of minimizing transonic and supersonic wave drag. At subsonic Mach numbers, however, the disadvantages are dominant. They include high induced drag (due to small wing span or low aspect ratio), high angles of attack for maximum lift, and reduced effectiveness of trailing-edge flaps. The straight-wing airplane does not have these disadvantages. For an airplane which is designed to be multimission, for example, subsonic cruise and supersonic cruise, it would be advantageous to combine a straight wing and swept wing design. This is the logic for the variable sweep or swing-wing. Although not necessarily equal to the optimum configurations in their respective speed regimes, it is evident that an airplane with a swing-wing capability can in a multimissioned role, over the total speed regime, be better than the other airplanes individually.

The swing-wing design attempts to exploit the high lift characteristics of a primarily straight wing with the ability of the sweepback wing to enable high speeds. During landing and takeoff, the wing swings into an almost straight position. During cruise, the wing swings into a sweepback position. There is a price to pay with this design, however, and that is weight. The hinges that enable the wings to swing are very heavy. One major drawback of the swing-wing airplane is the added weight and complexity of the sweep mechanisms.

Swept wings decreased controllability and combat load at takeoff, unless the wings could be pivoted forward during takeoff and landing and swept back during flight. Test articles from wartime German experiments had pointed the way, and the Bell X-5 provided additional data during the early 1950s. The British also had a variable-sweep concept plane called the Swallow, which underwent extensive testing at Langley. The NASA contribution in this development included variable in-flight sweeping of the wings and the decision to locate the pivot points outboard on the wings rather than pivot the wings on the centerline, solving a serious instability problem. All of this eventually led to the TFX program, which became the F-111. It was a long and controversial program but the eventual success of the variable geometry wing on the F-111 and the Navy's Grumman F-14 Tomcat owed much to NASA experimental work.

Boeing began small-scale studies of SST designs in 1952, and in 1958 established a permanent research committee which grew to a $1 million effort by 1960. The committed evaluated a number of alternative designs, all under the name Model 733. Most of the configurations were large delta wing planforms. In 1959 one swing-wing design was evaluated based on Boeing's efforts in the TFX project. In 1960 an evaluation of a baseline 150-seat aircraft for trans Atlantic routes incidated that the swing-wing version promised to be considerably better than the delta wing designs.

NASA did considerable work, starting in 1959, on basic configurations for the SST. In 1959, a delegation from NASA Langley briefed E. R. Quesada, head of the FAA, on the technical feasibility of a supersonic transport (SST). The NASA group advocated a variable geometry wing and an advanced, fan-jet propulsion system. The briefing, later published as NASA Technical Note D-423, "The Supersonic Transport: --A Technical Summary," analyzed structures, noise, runways and braking, traffic control, and other issues related to SST operations on a regular basis. An SST, the report concluded, was entirely feasible. The FAA concurred, and within a year, a joint program with NASA had allocated contracts for engineering component development. Eventually, the availability of advanced Air Force aircraft provided the opportunity to conduct flight experiments as well. There evolved four basic types of layout which were studied further by private industry. Lockheed chose to go with a fixed-wing delta design; whereas, Boeing initially chose a swing-wing design.

Many of the airplane-configuration studies undertaken by NASA during the 1959-62 period were concerned with the design of the North American B-70 Mach 3 bombing airplane and with preliminary designs for a commercial supersonic transport. The SST had become of interest as a national development project and it was felt to be the responsibility of NASA to take the lead in determining the most promising general configurations from which a successful SST might be developed.

Toward this end, the NASA Langley Research Center sponsored what was known as the SCAT (supersonic commercial air transport) program to investigate four basic SST configurations designated SCAT 4, 15, 16, and 17. SCAT designs 4 and 15 were quickly disposed of, leaving SCAT 16, having a wing with controllable sweep, and SCAT 17, having a canard configuration with a fixed delta wing. These two design arrangements were studied intensively both by NASA, in its wind tunnels, and by certain major aircraft companies. Ames, because of its long-standing interest in delta-wing and canard configurations, gave most of its attention to SCAT 17; whereas Langley, for similar reasons, devoted most of its efforts to SCAT 16. The Boeing and the Lockheed aircraft companies were later awarded Government contracts for SST designs based essentially on the SCAT 16 and the SCAT 17 configurations.

Boeing's design, submitted to the FAA on 15 January 1964, was nearly identical to the swing-wing Model 733 studied in 1960. Known officially as the Model 733-197, it was also referred as either the 1966 Model and the Model 2707. This latter name became the best known to the public, while internally Boeing continued to use 733 designation. The design was surprisingly similar to the B-1 Lancer strategic bomber, though the engines were mounted in individual nacelles rather than the twin pods on the Lancer.

The size of the SST design grew to meet airline payload requirements. The new SST was intended to carry 250 passengers, more than twice as many as the Concorde, and to fly at Mach 2.7, versus the Mach 2.2 (1,400 mph) of the Concorde. It was required to have a trans-Atlantic range of 4,000 miles. The higher speed required the aircraft be built of either stainless steel or titanium. At speeds above Mach 2.2 atmospheric friction would cause the aluminium used in the Concorde or Tu-144 to loose structural integrity.

A downselect resulted in North American and Curtiss-Wright being dropped from the program, with both Boeing and Lockheed asked to offer designs using either of the remaining engine designs and able to meet more demanding FAA requirements. In November 1964 a second design review was held. Boeing had enlarged the original design into a 250-seat model, the Model 733-290.

Both Boeing and Lockheed prepared more detailed designs for final source selection in 1966. Boeing's configuration was the 300-seat Model 733-390. Problems with this wing-tail configuration were exhaust scrubbing and acoustic noise/fatigue on the passenger cabin and aft fuselage; and pitch-up in both swept and unswept wing positions. Shortly before the end of the final competition, concerns about the Boeing design led to a drastic redesign. In the original configuration, the engines were to have been under the fixed portion of the swing-wing, but in response to design concerns -- including fears about what the jet exhaust might do to the tail (some suggested it might burn off!) -- the wings, tail and engine arrangement were drastically reworked for the competition deadline.


The winning 1967 Boeing SST proposal 2707-100, while still a variable sweep wing design, reverted to a high sweep, low aspect ratio planform when the wing, in the most aft swept position, was integrated and locked onto the horizontal stabilizer. The four turbojets were mounted beneath the horizontal stabilizer exhausting behind the trailing edge. The resulting high-speed configuration was a classic slender delta with a long, overhanging forebody.

Around the time it submitted its entry, Boeing made a few more modifications to the 2707 design. The modified 2707-100 design reduced minimum wing sweep from 30 to 20 degrees for better takeoff performance and widened the fuselage to accomodate 2-3-2 instead of 3-3 seating for much of the cabin's length.

The model 2707 in addition to swing-wings, featured a distinctive double-hinged "droop snoot" nose to allow optimal visibility during takeoffs and landings. The main reason for the double-jointed nose section was for ground clearance with the nose drooped. Major design changes were incorporated into the Boeing 2707-100 design. The supersonic cruise lift-drag ratio increased from 6.75 to 8.2, and the engines were moved farther back to alleviate the exhaust impinging on the rear tail surfaces. The four engines had been moved to underneath an enlarged tailplane. When the wings were in the swept-back position, they merged with the tailplane into a delta-wing planform.

Full-scale mock-ups of both the Boeing 2707-100 and Lockheed L-2000 were presented in September 1966. Lockheed's L-2000 was assessed as offering lower risk and simpler production, but providing slightly lower performance and somewhat higher noise levels. On 31 December 1966 Boeing was announced as the winner, powered by the General Electric GE4/J5 engines. Given the FAA's mandate to produce an advanced design, the decision was not surprising.

The Model 2707-100 was an advanced aircraft, even at subsonic speeds. The modified 2707-100 SST was initially intended to carry 277 passengers (a very large number at the time), fly at Mach 3, and have a range of 4,000 miles. The mock-up was fitted with 277 seats (30 first-class and 247 tourist). The impression on entering the cabin was that the so-called "narrow" part of the fuselage was noticably wider (about 4 ft or 1.22 m) than any contemporary jet transport. The cabin length was interrupted by two galley/toilet areas. Seating was to be seven abreast, two seats each side with three in the center, and two aisles. One of the earliest wide body designs, the 2-3-2 row seating was in a fuselage that was wider than aircraft then in service. The mock-up included both overhead storage with restraining nets for small items, along with large drop-down bins between sections of the aircraft. Wardrobe racks, galley tray containers and bar units could be removed from stowed positions and wheeled up and down aisles. Overhead luggage racks included restrainers, and were capable of housing items which usually had to be stowed under passengers' feet. The main 247-seat tourist-class cabin featured an entertainment system consisting of retractable televisions sets in the overhead storage between every 6th row. In the 30-seat first-class area every pair of seats included smaller televisions in a console between the seats. Windows were only 6" due to the high altitudes the aircraft flew at maximizing the pressure on them, but the internal pane was 12" to give an illusion of size.

The 2707 used four General Electric GE4/J5P turbojets, each of 63,200 lb. st (28677 kgp) each, with augmentation, were similar to the YJ93-GE-3 used on the XB-70. The General Electric YJ93-GE-3 engine originally used in the XB-70 supersonic bomber was mainly a subsonic engine with an impressive speed capability with afterburner lit, but with resulting poor fuel economy. The YJ93 engine is an afterburning turbojet designed for cruise at a Mach 3 flight speed, somewhat faster than the Mach 2.7 speed of the 2707. With the GE4/J5P, an afterburner would be used only for takeoff and transonic acceleration, with dry thrust sufficing for Mach 2.7 cruise.

The GE4 was 63,200 lb. thrust, single rotor augmented turbojet. The compressor is of medium pressure ratio with 9 axial stages. Two gangs of variable stators, front and rear, regulate air flow in subsonic and supersonic operation. The compressor is driven by a two-stage, air-cooled turbine. Advanced cooling techniques allow continuous operation of turbine inlet temperature in excess of 2O0OF. A conventional three bearing support system with three main structural frames is used to permit close clearance control in both the compressor and turbine. An annular combustor system was selected to achieve more uniform temperature distribution to the tu bine and insensitivity to inlet distortion. Th edesign also eliminates many of the troublesome problem makers, such as cross tubes and transition sections, of current cannular systems. Thrust augmentation is provided by a conventional afterburner. The exhaust system features a variable area, guided expansion system designed for high performance over a wide range of pressure ratios. The thrust reverser is integrated in to the primary nozzle for minimum weight and reduced system complexity.

The engines for Boeing's 2707 was required to operate in an environment far more hostile than had existed for any engine of commercial flights. The expected mission saw more than 70% of the engine flying time was to be spent in the hostile environment of supersonic flight. Requirements for safety, reliability, utilization, maintainability, performance and life, therefore, exceeded considerably the requirements of previous commercial equipment. When engineers looked at all of these factors the challenge of the assignment was obvious. Advanced technology was utilized in all areas to assure a dependable economical engine for this aircraft.

The inlet duct design for the supersonic propulsion system was substantially more complex than a duct for a subsonic engine which basically has no more than an aerodynamically-clean duct to provide for the passage of airflow from free stream conditions to the engine. The supersonic inlet at Mach 2.7 operates in an environment where the theoretical ram pressure ratio from free stream is 23:1, compared to less than 2:l at subsonic speeds. For the supersonic inlet to efficiently convert the available kinetic energy of the air to pressure and suhsonic velocities which the engine can accept, an internal-external compression passage with close control of the internal contraction ratio (ratio of throat area to flow area at the inlet lip) is required. The internal contraction ratio of the SST axially symmetric inlet is adjusted by a variable centerbody witbin a converging internal cowl wall. Both the centerbody and cowl wall contours are dictated by the internal shock pattern at Mach 2.7 cruise.

Boeing projected construction of the prototypes would begin in early 1967 and the first flight could be made in early 1970. Production aircraft could start being built in early 1969, with flight testing in late 1972 and certification by mid-1974.


During the prototype phase Boeing encounted insurmountable weight problems due to the swing-wing mechanism. Continued concerns about the aircraft's stability and load-carrying ability led to this second major revision of the design. The model 2707-200 lengthened "The Monster" and added small canards to its front [the word means "duck" in French]. A canard is a lifting airfoil located in the front portion of an airplane that compensates for the absence of a tail-mounted horizontal stabilizer. However, over time it became apparent that the weight of the swing-wings and all its related slats and control surfaces would make the plane rather heavy.

For a typical canard configuration, the canard tail surfaces are mounted forward on the fuselage and are designed to stall before the aft-mounted main wing. The mechanism of canard stall (and the associated loss of canard lift and the effectiveness of canard-mounted elevators) results in an inherent limiting of angle of attack to values lower than that required to stall the main wing.

Thc basic task of the 2707-200 Baeing SST was to increase productivity, in terms of payload/miles, over that of subsonic transports while retaining or improving upon their economy and safety levels. Some of the difference between the Boeing SST and subsonic transports, apart from the larger speed range up to M 2.7, as the larger gross weight and mass moments of inertia (the pitch moment of inertia was approximately ten times that of the 707-320). It also experienced a large range of environmental temperatures with maximums over 400F due to aerodynamic heating. The wings featured reduced thickness ratio fixed and moveable flight surfaces dictated by aerodynamic drag aerodynamic performance. The variable sweep wing geometry provided advantages in aerodynamic performance.

The 2707-200, while performing throughout a much larger flight envelope with no increase in required flight crew skill levels, required a Flight Control System with reliability equal to or exceeding that of existing subsonic transports. Relinbility for safc operation is assigned a top priority, while reliability to assure scheduled operation is recognizcd as an economic neccssity. Redundancy was employed in the form of multiple hydraulic power systems, surface actuators, mechanical load paths, and electronic subsystems. A comparison of control surface hinge moments with those of subsonic jets dictates a fully powered actuation system with no direct pilot manual reversion. Long mechanical control system runs required the incorporation of pilot assist servos to obtain surface resolution. In ordcr to reduce flight crew work load and attain the increased system performance requirements, utilization of electronic control subsystems was increased. An electric command subsystem was combincd with a mechnnical system for optimum performance and reliability. Hydraulic and electronic systems were designcd to remain operational following the failure of one system and to provide safe flight following failure of two systems.

The flight control system must provide airplane control and handling qualities as good as those of subsonic jets with no increase in flight crew skill levels. Although an increase in automatic and electronic flight control modes was indicated to minimize flight crew work load and supplement their inherent capabilities, the pilot would have the capability of overriding all such modes. This was to be accomplished as much as possible by utilizing his natural reactions.

Trailing-edge flight control surfaces were utilized on the fixed horizontal stabilizer and on the moveable wings for roll and pitch control. Lift reducing spoilers were also used to supplement roll control. A multiple segmented rudder was employed for directional control. With moveable wings swept aft for supersonic flight, the resulting mating with the horizontal stabiliher produced one continuous airfoil section. When the wings were swept forward for subsonic cruise and low speed landing and takeoff conditions, trailing-edge flaps and leading-edge slats were extended.

One problem associated with the SST is the tendency of the nose to pitch down as it flies from subsonic to supersonic flight. The swing-wing can maintain the airplane balance and counteract the pitch-down motion. Lockheed needed to install canards (small wings placed toward the airplane nose to counteract pitch down. Eventually, the Lockheed L2000 design had used a double-delta configuration and the canards were no longer needed. This design proved to have many exciting aerodynamic advantages. The forward delta begins to generate lift supersonically (negating pitch down). At low speeds the vortices trailing from the leading edge of the double delta . This means that many flaps and slats could be reduced or done away with entirely and a simpler wing design was provided. In landing, the double delta experiences a ground-cushion effect which allows for lower landing speeds. This is important since three-quarters of the airplane accidents occur in take-off and landing. The British-French Concorde and the Russian TU-144 used a variation of the double delta wing called the ogee wing. It, too, used the vortex-lift concept for improvement in low-speed subsonic flight.

Supersonic transports would land nose-high, with the flight deck 45 feet above the runway and more than 50 feet forward of the landing gear. In that position, the pilots have no view of the runway ahead of them. In the first generation supersonic transports - the 2702, Concorde and the TU-144 - the forward vision problem was solved by use of a mechanism that lowers-or "droops"-the forward part of the nose section for takeoffs and landings and thereby affords a clear view forward.

Like the Concorde, the 2707-200 SST had a variable nose geometry to improve flight deck forward views on approach. Boeing used a double-hinge, with the section forward of the cockpit angling down but the nose cone maintaining a similar axis to that of the fuselage. With the nose raised, minimum ground clearance was 8 ft. 9 in (2.67 m), reducing to only 4 feet (1.22 m) with it lowered.

There were overwhelming technical problems associated with the variable sweep wing design. These problems included aeroelastic effects due to the long fuselage, the need for a canard to meet takeoff rotation requirements, low values of lift-to-drag ratio for loiter due to outboard panel stall, and main landing gear placement in relation to engine location.

The hinge mechanism for the swing wings presented the greatest problem. For maximum effectiveness swing wings must have their pivots as close as possible to the centerline of the aircraft, so that the greatest benefit of increased wing span can be so achieved. But this interfered with the undercarriage and the positions of the engines. And the hinge mechanism was so heavy that it negated the advantage that gave the swing-wing configuration.

On 15 January 1968 an FAA technical team began a review of modifications made by Boeing to its SST prototype design (variable-sweep-wing model 2707-200). The team found that these changes, by increasing the aircraft's weight, had resulted in a poor weight-payload ratio. This overweight factor limited range and payload to such an extent that the prototype's calculated performance fell well below the specifications for the Phase III contract. With a full payload, the 2707-200 had a range of only 2,775 statute miles. An amendment to the Phase III contract, dated Mar 29, 1968, required Boeing to submit to FAA by 15 January 1969, a fully substantiated design capable of meeting the Phase III contract criteria for the prototype airplane.


Despite the advantages previously quoted for a swing-wing concept, technological advances in construction did not appear in time. Because of the swing-wing mechanisms and beefed-up structure due to engine placement, incurable problems in reduction of payload resulted. Boeing had no recourse but to adopt a fixed-wing concept-the B2707-300. The tailed delta wing was similar to the design proposed by Lockheed. On 21 October 1968 the Boeing Company formally announced it had abandoned its variable-sweep-wing design for the U.S. supersonic transport (SST) in favor of a conventional fixed-wing. The company's engineers had never been able to overcome the weight penalties imposed by the variable-sweep wing design. Boeing would submit the new design to FAA for approval in January 1969.

The proposed Boeing 2707-300 SST was designed for 290 passengers. It had a 69,000-pound payload with four turbofan engines mounted about the center section of the wings, which carry JP-4 propellant. The vehicle had a reference lengtb of 315 feet, a wingspan of 126.8 feet, and a total gross weight of 640,000 pounds.

Use of stability augmentation methods during preliminary design led to a 150 inch reduction in fuselage length for the Boeing 2707-300 SST. The shortened fuselage also led to reduced vertical tail size and gear length, with a weight savings of 6,000 lbs and a range increase of 225 nautical miles. The weight savings reported by Boeing came at the expense of an increase in control system development cost, however. The total cost of the Boeing SST flight and avionics systems were estimated to be double that of the Boeing 747. As a result, there was an assumption that the increased flight control system design complexity and cost (risk) was balanced by the performance improvements in the new design.

Theoretical analyses and wind tunnel tests of a low-T speed flutter model (1/20 scale) of the B-2707-300 airplane, were conducted under the supersonic transport (SST) Follow-on Program-Phase XI. Activation of the two ailerons was attempted for purpose of flutter suppression and activation of the horizontal stabilizer (with geared elevator) was attempted for purpose of rigid-body stability augmentation. For flutter suppression, the activated system should be located as near the tip of the wing as possible. The outboard aileron measures 13.4% of the wing semi-span and 26% of the wing chord. Its mid-span line is located around 72% of the wing semi-span.

Two constraints of the airplane made a flutter-free design unusually difficult: 1) the relatively low payload/total weight ratio made additional structural weight or mass balance particularly distasteful, and 2) any arrangement of lifting surface planforms, thickness, or major mass relocation (e.g., nacelles) degraded the delicate cruise economy or c.g. balance. Because of this flutter dilemma, considerable efforts were directed towards the development of an active flutter suppression system with the objective of improving the flutter speeds of the SST airplane. The developed flutter suppression system yielded only minor improvements in flutter speeds (9.4% increase with activated inboard ailerons, 3.2% increase with activated outboard ailerons, and 11.3% increase with activated inboard and outboard ailerons).

Work began on a full-sized mockup and two prototypes in September 1969, now two years behind schedule. A promotional film claimed that airlines would soon pay back the federal investment in the project, and it was projected that SSTs would dominate the skies with jumbo jets being only a passing intermediate fad.

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Page last modified: 07-07-2011 02:27:23 ZULU