Category AERONAUTICS

"R. T.” Jones was a brilliant, flight-obsessed, and largely self-taught fluid dynamicist, having dropped out of the University of Missouri to join a flying circus, then working as a designer for Nicholas-Beazley, a small Missouri aircraft company. When the Great Depression collapsed the firm, his father used political connections as Chairman of the local Democratic Party to secure Jones a job running elevators in the U. S. Capitol. In his spare time and evenings, he studied mathematics and aerodynamics with Albert Zahm, the aeronautics Chair at the Library of Congress, and with Max Munk at Catholic University. Despite his lack of a formal engineering degree, through the efforts of Representative David Lewis (a homespun Maryland progressive with a strong interest in self-improvement who had taken math instruction from the young elevator operator), Jones received a temporary appointment as a "sci­entific aide” to the NACA. There, he quickly proved such a gifted and insightful researcher that he soon secured a coveted permanent posi­tion at Langley, consorting with the likes of John Stack, Eastman Jacobs, and Theodore Theodorsen.[17]

As he considered Griswold’s "glomb,” Jones recognized that its extremely low aspect ratio shape (that is, a shape having a very long

wing root in relation to its total wingspan) could not be adequately ana­lyzed using conventional Prandtl-rooted "lifting line” theory. Instead, Jones drew on the work of his mentor Munk, using papers that Munk had written on the flow of air around inclined airship hulls and swept wings, and one by the Guggenheim Aeronautical Laboratory’s Hsue-shen Tsien, a von Karman associate at the California Institute of Technology (Caltech), on airflow around inclined bodies of revolution. He analyzed it using linear equations governing two-dimensional incompressible flow, con­sidering his results of little practical value, recalling three decades later, "I thought, well, this is so crude, nobody would be interested. So I just hid it in my desk.”[18]

But it sparked his curiosity, and in January 1945, by which time he was busy thinking about nonlinear compressible flows, he had a rev­elation: the equations he had developed months earlier for the glomb analysis could be applied to a low aspect triangular wing operating in supersonic flow, one whose wing-leading edges were so sharply swept as to place them within the shock cone formed around the vehicle and hence operating in subsonic flow. In these conditions, the wing was essentially "fooled” into behaving as if it were operating at a much lower Mach number. As Jones recalled, "It finally dawned on me that the slen­der wing theory would hold for compressible flow and even at supersonic speed if it were near the center of the Mach cone. So, I immediately got the paper out and I added the compressible flow parts to it, which was really the important part, and then I wondered well, why is it that this slender wing doesn’t have an effect on compressibility? Then I realized that it was because the obliquity of the edge and that this is the sim­ple sweep theory and would work in spite of the compressibility effect. So, I wrote a paper which incorporated the slender wing theory and also sweep theory.”[19] Jones then moved from considering a slender triangular delta [Д] to the sharply sweptback wing [л], the reverse of

Germany, where the high-speed swept wing had preceded, not followed, the delta.[20]

Jones’s delta and swept wing utilized, for their time, very thin airfoil sec­tions, ones typical of supersonic aircraft to come. In contrast, German swept and delta wing developer Alexander Lippisch had employed much thicker sections that proved unsuitable for transonic flight. His tailless rocket – propelled swept wing Me 163 Komet ("Comet”) interceptor, for example, essentially became uncontrollable at speeds slightly above Mach 0.82 thanks to stability changes induced by shock wave formation on its relatively thick wing. His design for a rocket-boosted, ramjet-powered delta fighter, the P 13, had such thick wing and tail sections—the pilot actually sat within the leading edge of the vertical fin—that it could never have achieved its desired transonic performance. As discussed subsequently, postwar NACA tests of a captured glider configuration of this design, the DFS DM-1, confirmed that transonic delta wings should be far thinner, with sharper leading edges. As a consequence, NACA researchers rejected the Lippisch approach, and, though some of them tried extrapolations of his designs (but with lower thickness-chord ratios and sharper leading edges), the NACA (and industry as well) adapted instead the thin slender delta, a la Jones.[21]

In February 1945, Jones showed his notes on sweep to Jean Roche, the Army Air Forces technical liaison at Langley, and informed others as well, including Maj. Ezra Kotcher of the AAF’s Air Technical Service

Command, and NACA colleagues Arthur Kantrowitz and Hartley A. Soule.[22] Kotcher passed it along to von Karman and Tsien—then work­ing as scientific advisers to Gen. Henry H. "Hap” Arnold, the Army Air Forces’ Chief of Staff—and Soule and Kantrowitz urged Jones to inform the Agency’s Director of Research, George W. Lewis, of his discovery.[23] Accordingly, on March 5, 1945, Jones informed Lewis, "I have recently made a theoretical analysis which indicates that a V-shaped wing travel­ling point foremost would be less affected by compressibility than other planforms. In fact, if the angle of the V is kept small relative to the Mach angle [the angle of the shockwave], the lift and center of pressure remain the same at speeds both above and below the speed of sound.”[24] Jones subsequently undertook tests in the Langley 9-inch supersonic tunnel of a small, 4-inch-long daggerlike sheet-steel triangular wing with rounded leading edges and a span of only 1.5 inches, tests that complemented other trials at Aberdeen, MD, arranged by von Karman and Tsien.

The Langley tests, through the transonic region and up to Mach 1.75, confirmed his expectations, and Jones published his first test results May 11, 1945, noting, "The lift distribution of a pointed airfoil travelling point-foremost is relatively unaffected by the compressibility of the air below or above the speed of sound.”[25] This was almost 2 weeks before Lippisch informed von Karman, then leading an AAF European study team, of his high-speed delta concepts (during a technical intelligence interrogation at St. Germain, France, on May 23), not quite a month before von Karman assistant Clark Millikan visited the Messerschmitt advanced projects group at Oberammergau on June 9-10 and inter­rogated Waldemar Voigt about his swept wing fighter concepts, and well over a month before Millikan journeyed to Volkenrode to inter-

rogate German swept wing inventor Adolf Busemann, on June 20-21.[26]

Langley’s peer reviewers and senior Agency official Theodore Theodorsen did not immediately accept Jones’s assumption that a unified slender wing theory could apply to both compressible and incompressible flows and even questioned the evidence of sweep’s benefits. Fortunately, Jones was greatly assisted in confounding skeptics by the timely results of NACA tunnel tests and falling body experiments, which left little doubt that sweep worked. As well, an associate of Jones made a most helpful discov­ery: locating a 1942 British translation of Busemann’s 1935 paper. Evidence of an enemy’s interest coincident with one’s own work always heightens its perceived value, and undoubtedly, the Busemann paper, however dated, now strongly bolstered Jones’s case. When it became time to assemble a bibliography for his swept wing report, Jones added Busemann’s paper and other German sources by Albert Betz, H. G. Kussner, Ludwig Prandtl, and Hermann Schlichting, though it is unclear whether this reflected a collegial respect across the chasm of war or simply a shrewd apprecia­tion of their persuasive value.[27]

Langley released his report in late June 1945.[28] In it, Jones noted: "the attachment of plane waves to the airfoil at near-sonic or supersonic speeds (Ackeret theory) may be avoided and the pressure drag may be reduced by the use of planforms in which the angle of sweepback is greater than the Mach angle. The analysis indicates that for aerodynamic efficiency, wings designed for flight at supersonic speeds should be swept back at an

angle greater than the Mach angle and the angle of sweepback should be such that the component of velocity normal to the leading edge is less than the critical speed of the airfoil sections. This principle may also be applied to wings designed for subsonic speeds near the speed of sound, for which the induced velocities resulting from the thickness might otherwise be suf­ficiently great to cause shock waves.”[29] Such marked the effective birth of
the high-speed swept wing airplane in the United States, as his report weeks earlier had marked the birth of the American high-speed delta.

By the time Jones’s report appeared, Germany’s aeronautical estab­lishment was already under the microscope of Allied technical intelli­gence, whose teams swiftly focused on its intensive investment in swept wing aerodynamics for its missiles and aircraft. Replicating reaction to the earlier "discovery” of "Gottingen aerodynamics” after the First World War, the post-Second World War influence of German example and practice was even more profound. Indeed, it affected the entire post­war course of European, Soviet, and American high-speed aerodynamic research, development, test, evaluation, and acquisition. In the increas­ingly tense national security environment of the burgeoning Cold War, the national intelligence services of the various advanced aeronauti­cal nations understandably maintained very active technical collection efforts to learn what they did not already know.[30]

The NACA so rapidly focused its attention on swept planforms that, within 2 years of the end of the Second World War, George Gray, author of a popular yet surprisingly detailed study of the Agency, could already write: "Just how far the sweepback principle can be applied with result­ing advantage is a question. . . . At about 90 percent of the speed of sound both sweepback and low aspect ratio begin to be of value, and wings that combine the two features seem to offer a promising choice. At about Mach number 1.50, a sweepback of 60 degrees seems neces­sary to escape the backward flare of the Mach angle. . . . At Mach num­ber 2.00, the angle is so acute that it is impossible to avoid it and still preserve the wings. It may be that designers preparing for flight at this speed will return to wings of low angles of sweep, and place their main dependence for drag reduction on thinning the profiles, lowering the aspect ratio, and sharpening the edges of wings.”[31] By 1950, this grow­
ing confidence in the old-new swept planform had resulted in transonic and supersonic research airplanes, a variety of military prototypes, and two operational jet fighters that would shortly clash over North Korea: the American F-86 Sabre (first flight in October 1947) and, in the Soviet Union, the MiG-15 (first flight in December 1947).[32]

Swept wing aircraft, for all their high-speed advantages, posed daunt­ing stability, control, and handling qualities challenges. Foremost of these was pitch-up at low and high speeds, resulting from deteriorating longitudinal stability.[33] A swept wing airplane’s lateral-directional sta­bility was compromised as well by so-called "dihedral effects.” Swept wing aircraft with excessive dihedral experienced pronounced combined rolling and yawing "Dutch roll” motions, which would be unacceptable on both production civil and military designs.[34] Such motions would induce airsickness in passengers on large aircraft and, on bomber, fighter, and attack aircraft, prevent accurate tracking of a maneuvering target or accurate bomb release. (Indeed, it was largely because of this kind of behavior that the U. S. Air Force did not proceed with production of Northrop’s YB-49 flying wing jet bomber.) Adverse yaw posed another problem. At higher speeds, as a swept wing plane rolled from aileron

deflection, it experienced higher drag and loss of lift involving the low­ered wing, generating a tendency of the airplane to turn (reverse) into the direction of the raised wing, effectively doing the opposite of what the pilot intended. Adverse yaw could be caused by aeroelastic effects as well. That swept wing aircraft would possess behavior characteris­tics significantly different than conventional straight wing designs did not come as a surprise to the NACA or other aerodynamic researchers in America and overseas. But all recognized the need to complement theory and ground-test methodologies with flight research.

The peculiarities of swept wing aircraft, at a time when early jet air­craft lacked the power-to-weight advantages of later designs, could—and often did—prove fatal. For example, Boeing designed the B-47, America’s first large swept wing aircraft, with pod-mounted engines and a broad, highly tapered, thin swept wing. During flight-testing at higher speeds, test pilots found aileron input to roll the aircraft would twist the wing, the aileron effectively acting as a trim-tab does on a control surface. The twisted wing would overcome the rolling moment produced by the aile­ron, rolling the aircraft in the opposite direction. Aeroelastic structural divergence caused several accidents of the B-47 during its flight-testing and service introduction, forcing the Air Force to limit its permissible airspeed to 425 knots, as high as it could be safely flown if roll reversal were to be avoided. As a result, Boeing built its successors, the XB-52 and the Model 367-80 (prototype for the KC-135 family and inspiration for the civil 707), with much thicker wing roots and structures that were torsion resistant but that could still flex vertically to absorb structural loads and gust-induced loads during flight.[35]

But the most serious swept wing problem in the early jet era was pitch – up, a condition affecting both low – and high-speed flight, reflecting stall onset either from decreasing speed (low-speed pitch-up) or from trim changes during high-speed flight, particularly during accelerated
maneuvers, such as "wind-up” turns that rapidly increased g-loading and angle of attack. Pitch-up occurred at the breakpoint in a lift curve, immediately beyond the peak point where the airplane’s wing was oper­ating at its highest lift-producing angle of attack, with its lift coefficient at maximum value. At the breakpoint, the wing would begin stalling, with flow separation from the airfoil, breaking the circulatory flow pat­tern around the wing. In ideal circumstances with a straight wing air­craft, the change in lift would occur simultaneously spanwise across the wing and would typically trigger a nose drop. But in a swept wing aircraft, the stall would first begin at the tips and progress inward, the center of lift shifting forward. As the plane’s longitudinal (nose-up nose – down) stability decreased, the shifting center of lift would abruptly rotate the nose upward (hence the use of the expression "pitch-up”), even at a rate of onset beyond the capabilities of its elevator control surfaces to correct. As well, of course, since the ailerons that governed lateral con­trol (roll control) were typically located outboard on a wing, a swept wing airplane could lose its lateral control authority precisely at a point when the pilot needed as much control capability and reserve as possi­ble. Because stall onset is not always triggered uniformly, a swept wing airplane nearing the pitch-up point could experience sudden loss of lift on one wing, inducing abrupt rolling motions (called "wing dropping”), complicating its already dangerous low-speed behavior.

There was, of course, the possibility of overcoming such problems by sweeping a wing forward, not aft. A forward-swept wing (FSW) had both desirable high – and low-speed aerodynamic characteristics. Since the spanwise flow would run from the tips to the fuselage, the outer portions of the wing would stall last, thus preserving lateral control. As well, it would have more desirable pitching characteristics. Already, in the midst of the Second World War, the Germans had flown an exper­imental bomber, the Junkers Ju 287, featuring a forward swept wing, and a number of aircraft and missile projects were forecast for such planforms as well. The forward swept wing, and combined-sweep "M,” "W,” and even "X” planforms, received a great deal of postwar atten­tion, both in America and abroad. Researchers at Langley modified wind tunnel and configuration models of both the XS-1 and D-558 to employ forward-swept wing planforms, and tested conceptual planforms with both aft and forward sweep to develop comparison data. But while the FSW undoubtedly had better low-speed behavior, it had higher profile drag and posed difficult structural problems for designers. In

the precomposite structure era, an FSW had to be necessarily heavier than an aft-swept wing to avoid aeroelastic flexing that could inhibit both good flight performance and even flight safety. Further, the struc­tural and weight limitations also limited the sweep angles that an FSW could then have; even as late as the 1960s, when Germany produced a business aircraft (the Hamburger Flugzeugbau HFB-320 Hansa Jet), it possessed only modest forward sweep and, though flown suc­cessfully and built in small numbers, was not a commercial success. It would take over three decades before the advent of computerized flight control, composite structures, and a more radical vision of for­ward sweep application would result in experimental planforms like the NASA-Rockwell HiMAT remotely piloted vehicle, the piloted Air Force – DARPA-NASA Grumman X-29 (and, in Russia, an X-29-like experimen­tal aircraft, the Sukhoi Su-37). Even so, and even though forward sweep would be applied to some weapon systems (for example, the AGM-129 stealthy cruise missile, where it contributed to its low radar reflectiv­ity), forward wing sweeping would remain the exception to "normative” aft-swept wing design practice.[36]

Pitch-up was profoundly dangerous. At low speeds in proximity to the ground, it could—and often did—trigger a disastrous departure and crash. The recognition of such problems had caused the U. S. Navy to procure two modified Bell P-63 Kingcobra fighters (designated L-39), which had their wing panels replaced with 35-degree swept wing sec­tions, and a fuselage extension to accommodate their now-changed cen­ter of lift. Not intended for high speeds, these two low-speed swept wing research aircraft were extensively flown by various contractor, Navy, and NACA research pilots to assess the basic behavior of the swept wing, with and without lift-and-control-augmenting devices such as wing slats and flaps. They quickly encountered its limitations. On one flight with the plane in "clean” (i. e., slat-free) configuration, Bell Company test pilot A. M. "Tex” Johnston gradually raised the nose of the plane while retarding power. After just "a slight tremor” indicating the onset of asymmetric tip stall, it "instantaneously rolled to an almost inverted

position.”[37] Grumman test pilot Corwin "Corky” Meyer recalled that while the L-39 was "docile” with leading edge slats, without them it "cavorted like a cat on catnip.”[38] The two L-39 aircraft furnished vital insight into the low-speed performance and limitations of swept wing aircraft, but they also clearly demonstrated that such aircraft could, in fact, be safely flown if their wings incorporated careful design and safety devices such as fixed leading edge slots or movable slats.[39]

In military aircraft, pitch-up could prevent a pilot from maneuvering effectively against a foe, could lead to loss of control of the airplane, and could result in such excessive airframe loadings that an airplane would break up. It was no respecter of designs, even outstanding ones such as North American’s evocative F-86 Sabre, generally considered the finest jet fighter of its time by both American and foreign test pilots. First flown in October 1947, the Sabre quickly became an internationally recognized symbol of aeronautical excellence and advancement. When British test pilot Roland Beamont, a distinguished Royal Air Force fighter ace of the Second World War, evaluated the Sabre at Muroc Dry Lake in May 1948

(a month after it had dived past the speed of sound, becoming the world’s first supersonic turbojet airplane), he likewise dived it through Mach 1, thus becoming the first supersonic British pilot. Afterward, he noted approvingly in his test report, "The P-86 is an outstanding aircraft.”[40] The Sabre’s reputation was such that British authorities (frustrated by the slow development pace of Albion’s own swept wing aircraft) tell­ingly referred to it simply as "That Aircraft.” Vickers-Supermarine test pilot David Morgan recalled, "No British fighter of the day could match the handling of the North American F-86.”[41] Indeed, designers from his company, frustrated by their slow progress turning the experimen­tal Swift into a decent airplane, even resorted to crude subterfuge in an effort to unlock the Sabre’s secrets. When a pair of Canadian pilots landed at the Supermarine plant in their Canadair-built Sabres, com­pany officials, with apparent generosity, laid on a fancy lunch, driving them off to a local hotel. While the visiting airmen dined and chatted with solicitous Supermarine representatives, another team of engineers "swarmed over the Sabres to study their construction,” marveling at "this splendid aircraft.”[42]

Yet however "splendid” "That Aircraft” might otherwise have been, the Sabre killed unwary pilots by the dozens in accidents triggered by its low-speed pitch-up tendencies. Apollo 11 astronaut Michael Collins recalled his introduction to the F-86 at Nellis Air Force Base as "a bru­tal process. . . . In the eleven weeks I was there, twenty-two people were killed. In retrospect it seems preposterous to endure such casualty rates without help from the enemy, but at the time the risk appeared perfectly acceptable. . . . I’m surprised to have survived. I have never felt quite so threatened since.”[43] In over a decade of tests with various Sabre variants

to improve their low-speed handling qualities, NACA Ames research­ers assessed a variety of technical "fixes.” The most beneficial was the combination of artificial feel (to give the pilot more reassuring higher maneuvering control forces during the approach-to-landing, combined with greater inherent stability than possible with a non-artificial-feel sys­tem), coupled with leading-edge suction to draw off the boundary layer airflow.[44] First evaluated on a test rig installed in the Ames 40-foot by 80-foot full-scale wind tunnel, the boundary layer control (BLC) exper­iment on the F-86 proved most valuable. Ames researchers concluded: "Leading edge boundary-layer control was most effective in providing a large reduction in both stalling speed and approach speed together with an increased margin of lift for flare and maneuvering during the [land­ing] approach,” an important point, particularly for swept wing naval aircraft, which had to be controllable down to a landing on the con­fined deck of an aircraft carrier.[45] The trials benefitted not only future swept wing studies but, more generally, studies of BLC applications for Vertical/Short Take-Off and Landing (V/STOL) aircraft systems as well.[46]

Nor were the Sabre’s high-speed pitch-up characteristics innocuous. The NACA flew extensive Sabre evaluations at its High­Speed Flight Research Station and at Ames to refine understanding of

its transonic pitch-up behavior, which test pilot A. Scott Crossfield recalled as "violent and dangerous.”[47] It could easily exceed its design load factors, sometimes pitching as high as 10 g. At 25,000 feet, in the very midst of its combat operating envelope (and at lift coefficients less than its maximum attainable lift) the Sabre’s pitch-up onset was so severe that g forces once momentarily "blacked out” the test pilot. Overall, after extensive Ames tests, the early slatted F-86A with a conventional fixed horizontal stabilizer and movable elevator was judged "unsatisfactory” by a group of highly experienced fighter test pilots, thanks to its "severe pitch – up tendencies.” The same group found the later slat-less "6-3” F-86F (so – called because its wing extended forward 6 inches at the root and 3 inches at the tip, a modification made by North American based on Korean war experience) had "moderate” pitch-up tendencies. Because of this, and because it had an adjustable (not fixed) horizontal stabilizer in addition to its elevator, the pilots judged the F-86F’s pitch-up behavior "unsatisfactory but acceptable.”[48]

Worse swept wing problems plagued the Sabre’s great adversary, the Soviet MiG-15. Unlike the Sabre, the MiG-15 had a less aerodynamically pleasing configuration, and its fixed horizontal stabilizer and elevator combination, located midway up the vertical fin, made it more suscep­tible to aerodynamic "blanketing” of the tail by the wing and, hence severe pitch-up problems, as well as limiting its transonic maneuver­ability (to the Sabre’s advantage). During the Korean war, Sabre pilots frequently saw MiG pilots eject from otherwise perfectly sound aircraft that had pitched up during turns, stalled, and entered flat, unrecoverable spins. Nearly five decades later, Soviet pilot Stepan Mikoyan (nephew of Anushavan "Artem” Mikoyan, cofounder of the MiG design bureau) con­ceded that high-speed accelerated stalls often triggered unrecoverable spins, leading to "a number of ejections and fatal accidents.”[49] Postwar American testing of a MiG-15 delivered by a defecting North Korean

pilot confirmed the MiG’s marked vulnerability to pitch-up-induced stalls and spins; indeed, the defector’s own instructor had been lost in one such accident. Not surprisingly, when Mikoyan produced the MiG – 17—the lineal successor to the MiG-15—it had a very different outer wing configuration giving it more benign behavior.[50]

Western European swept wing aircraft exhibited similar problems as their American and Soviet counterparts. For a brief while, influenced by the Messerschmitt Me 163 Komet and a variety of other German proj­ects, designers were enthralled with the swept wing tailless configura­tion, believing it could resolve both the challenges of high-speed flight and also furnish inherent stability.[51] Then, in September 1946, British test pilot Geoffrey de Havilland perished in an experimental tailless transonic research aircraft, the de Havilland D. H. 108 Swallow, when it began an undamped violently divergent longitudinal pitching oscilla­tion at Mach 0.875, breaking up over the Thames estuary and proving that the "sound barrier” could bite.[52] Subsequently the NACA evaluated the Northrop X-4, a generally similar American configuration. Tested at high altitude (and hence, at low dynamic pressure), the X-4 fortu­nately never "diverged” as violently as the ill-fated D. H. 108. Instead, as NACA pilot A. Scott Crossfield remembered, at Mach 0.88 "it broke

into a steady porpoising motion, like an automobile cushioning over a washboard road.”[53] Conventional tailed European swept wing designs followed the same steep learning curve as American ones. Britain’s Supermarine Swift, a much-touted design from the builder of the legendary Spitfire, had a "vicious” transonic pitch-up. By the time it entered service, it was years late, obsolescent, and useless for any other role save low-level tactical reconnaissance.[54]

Pitch-up afflicted a wide range of early transonic and supersonic jet fighters, and the NACA was fortunate in having an available research airplane that could study swept wing behavior across the transonic regime. This aircraft was the Douglas D-558-2 Skyrocket, "Phase II” of the larger D-558 research aircraft program, a Douglas Company ven­ture begun in 1945 and sponsored by the U. S. Navy and the NACA. The D-558 program had begun as a companion to the XS-1 effort and repre­sented a different design approach. Where the XS-1 was rocket powered, the D-558 Skystreak used a turbojet; where the XS-1 employed an ogi­val projectile shape with a midwing of 8-percent thickness-chord ratio, the D-558 used a constant-diameter tube wrapped around an axial-flow turbojet engine and a low wing of 10-percent thickness-chord ratio; and where the XS-1 was air launched, the D-558 took off from the ground as a conventional airplane. Both were straight wing designs, with their adjustable stabilizers and movable elevators placed midway up their ver­tical fins. All together, the Navy ordered six D-558 aircraft from the firm.[55]

Originally, swept wings had not featured in the D-558 program. Then the discovery by Douglas engineers of a plethora of German tech­nical reports (coupled with the work of Jones and others in the United States) caused the Navy, the NACA, and Douglas to modify the D-558
program.[56] The last three aircraft were completed as a new swept wing design. Initially, the planned modification seemed straightforward: replace the straight wing and tail surfaces with swept ones. In antic­ipation, Langley tested models of the D-558 with a variety of swept wings. But the possibility of giving the swept wing D-558 supersonic performance—something the D-558 straight wing lacked—resulted in a more radical redesign. Gone was the simple Pitot intake inlet. Instead, designer Edward "Ed” Heinemann and his team chose an ogival body shape resembling the XS-1. The new 35-degree slat-equipped swept wing was relocated to midfuselage position and given anhedral (droop), with the landing gear relocated into the fuselage. In contrast to the original single-engine D-558s, the new swept wing design featured both a 6,000- pound thrust rocket engine and a small turbojet. Thus recast, it received the designation D-558-2 and the name Skyrocket, to distinguish it from the straight wing Skystreak, itself redesignated D-558-1. The result was one of the most elegant and significant aircraft of all time.

The first D-558-2 flew in February 1948, though initial flight tests gave little hint of how remarkably versatile and successful it would prove. At max takeoff weight, it was so underpowered (and thus so slug­gish) that it needed four solid-fuel jettisonable assistance takeoff (JATO) rockets to help kick it aloft. Eventually, the Navy and the NACA would arrange to take the second and third Skyrockets and modify them for air launch from a modified PB2-1S (Navy B-29) Superfortress, dramatically improving both their safety and high-speed performance; fuel previously

spent climbing aloft could now be more profitably expended exploring the transonic and supersonic regimes. While the third aircraft retained its jet and rocket engine, the second had its jet engine removed and additional tanks for rocket propellant and oxidizer installed. Thus modified, the second aircraft reached Mach 2.01 in November 1953, flown by Scott Crossfield, the first piloted Mach 2 flight, having earlier attained an alti­tude of 83,235 feet, piloted by Lt. Col. Marion Carl, a noted Marine aviator. Eventually, the NACA received the first D-558-2 as well (which Douglas had employed for contractor testing). The Agency modified it as an all-rocket aircraft, though it only completed a single check flight before being retired.

Before all-rocket modification, the second Skyrocket introduced Agency pilots to the hazards of pitch-up. On August 8, 1949, during its seventh flight, pilot Robert Champine banked into a 4 g turn at Mach 0.6, and the Skyrocket violently pitched up, reaching 6 g. It responded rapidly to full-down elevator, and Champine made an uneventful (if pru­dently precautionary) landing. Thereafter, until returning the airplane to Douglas for all-rocket modification in 1950, the NACA flew extensive pitch-up investigations with it. In November, pilot John Griffith repli­cated the 4 g and Mach 0.6 pitch-up that Champine had experienced earlier. This time, however, he attempted to continue flying to more fully assess the Skyrocket’s behavior. Thus challenged, it snap-rolled on its back. After recovering, Griffith probed its low-speed behavior, gradually slowing, with flaps and gear extended and wing slats closed. At 14,000 feet and 130 mph, the Skyrocket pitched up, rolling into a spin, and los­ing 7,000 feet of altitude before its pilot could recover.[57] Clearly its ugly behavior did not match its alluring form.

Focused on extending the Skyrocket’s performance into the super­sonic regime by modifying the second aircraft as a pure rocket plane, the NACA turned to the third aircraft, which retained its jet engine as well as its rocket, for future pitch-up research. Air-launched, the jet-and- rocket Skyrocket had tremendous research productivity; it could accelerate

into the supersonic regime, above Mach 1.1, and its jet engine enabled it to "loiter” in the transonic region, making repeated data-gathering runs. Its comprehensive instrumentation package enabled assessment of loads, pressure distributions, and accelerations, evaluated against background data on flight conditions, aircraft attitude, and control surface position and forces. Between the end of 1950 and the fall of 1956, it completed 66 research flights on pitch-up and associated transonic phenomena, includ­ing the evaluation of the effects external wing stores—tanks and bomb shapes—had on aircraft performance. It evaluated a variety of proposed aerodynamic solutions and fixes to resolve the pitch-up problem, includ­ing various wing fence designs to "channel” airflow and inhibit the char­acteristic spanwise-flow (flow toward the wingtips) found with swept wing planforms, various combinations of slat and flap position, changes to lead­ing edge shape, and "sawtooth” leading edge extensions on its outer wing panels. All of this testing reinforced what engineers suspected, namely that no one overall technical fix existed that could resolve the pitch-up challenge. Rather, swept wing aircraft design was clearly situational, and, depending on the mission of the aircraft and its resulting design, combina­tions of approaches worked best, chief among them being low placement of

the horizontal tail, below the chord-line of the wing, coupled with provision of stability augmentation and pitch-damping flight control technology.[58]