SP-4302 Adventures in Research: A History of Ames Research Center 1940-1965

 

PART II : A NEW WORLD OF SPEED : 1946-1958

1954-1957

 

 

12

Research

 

PATTERN

 

[245] THE trend of Ames research toward the more scientific and the more fundamental continued during the period from 1954 to 1957. Many of the Laboratory's scientist-engineers were gaining worldwide reputations, and with increasing frequency they were invited to deliver papers before scientific societies at home and abroad. Among those who traveled abroad for this purpose were Max Heaslet (London, 1948), Jack Stalder (London, 1951), Ralph Huntsberger (The Hague, 1954), John Parsons (Stockholm, 1954), Harvey Allen (Paris, 1954), Walter Vincenti (Brussels, 1956), and Al Seiff (The Hague, 1957). As 1957 ended, John Spreiter was preparing a paper for delivery in Tokyo the following year.

Ames research was advancing along a broad front, but increasing emphasis was being given to the theoretical approach. Aerodynamicists were devoting much effort to the task of determining optimum configurations for supersonic and hypersonic aircraft. In flight research there was a continuing interest in automatic control and at the same time an accelerating trend in the use of flight simulators of both the airborne and the ground-based types. The later trend focused considerable attention on the interpretation of pilot opinion. Dynamic stability was receiving attention in all branches of the Laboratory during this period; also, efforts were being made to reduce the landing speed of airplanes by means of boundary-layer-control systems which utilized the surplus air-pumping capacity of the jet engines.

Perhaps the most notable feature of Ames research during this period, however, was the tremendous effort applied to related studies of boundary layer, skin friction, and aerodynamic heating. Although of a basic character, this research was aimed at a number of specific applications in which the Laboratory was particularly interested. These applications included hypersonic airplanes, man-carrying rocket-launched gliders called "boost-glide vehicles'' and ballistic vehicles. A variety of thermal-protection schemes for such craft was under consideration.

[246] This period was also noted for the initiation of research in the new Unitary Plan facility. The intended function of the new facility was to serve the development needs of the industry and the military and to free other facilities at Ames for research of a more fundamental character. Although at the end of World War II there had been a feeling that NACA should be relieved of some of its development-test burden, NACA's services to the military had proved much too valuable to be abandoned or even greatly diminished. Ames' contribution to the development-test effort since the end of the war was substantial and of great value; but, since it was often concerned with classified or proprietary matters, it is largely omitted from this history.

 

BASIC CONFIGURATlONS AND AIRFLOWS

 

wings. Important contributions to wing theory were made during this period by Max Heaslet, Harvard Lomax, R. T. Jones, Doris Cohen, John Spreiter, Alberta Alksne, and Milton Van Dyke. Work of the first four of these people appeared in volumes VI and VII of the monumental Princeton Series of books covering the general field of aeronautical science. The Princeton Series, organized under an editorial board headed by Dr. Theodore von Karman, presented material written by recognized American authorities. NACA and the Ames Laboratory were well represented in the group of authors. In volume VI, published in 1955, Heaslet and Lomax contributed a section dealing with supersonic and transonic small perturbation theory; and in volume VII, published in 1957, Jones and Cohen presented a very comprehensive review of the aerodynamics of wings at supersonic speeds. There were, it should be mentioned, many other Ames contributors to the Princeton Series. Among them were Harvey Allen, Joseph Spiegel, John Dimeff, Ben Beam, Jack Stalder, Al Seiff, and Alex Charters.

During these years, John Spreiter continued his work in the general field of transonic aerodynamic theory and, together with Alberta Alksne, turned out TR 1359 (ref. B-38), entitled "Thin Airfoil Theory Based on Approximate Solution of the Transonic Flow Equations." Also during this period, John and Alberta, together with A1 Sacks, were studying the vortex patterns shed by cruciform wing arrangements and the rollup of the vortex sheet behind such wings. This important work, having direct application to missiles with short wings, was published in TR 1296 and TR 1311.

Meanwhile, Milton Van Dyke, pursuing his studies of second-order flow theory, turned out a number of papers of which TR 1274, "second-order Subsonic Airfoil Theory Including Edge Effects," might be considered representative.

Experimental research on wings continued during these years but less of it was concerned with wings for what were at this time thought of as conventional transonic or supersonic airplanes. In the 1- by 3-foot tunnel, Walter Vincenti ran tests to confirm his earlier mentioned theory regarding the [247] flow over double-wedge airfoils operating just below the speed of shock attachment This work was reported in TN 3225 and TN 3522. It was as a result of his excellent work in developing and applying wing theory that Walter received the Rockefeller Public Service Award granting the privilege of a year of study at Cambridge.

One of the most outstanding results of experimental wing research at Ames was the development of conical camber as a means of reducing the drag-due-to-lift and thus of increasing the range of supersonic airplanes, particularly those equipped with delta wings. Conical camber was a form of camber which, for practical reasons, was restricted to the forward part of the wing. The portion of the wing involved in the camber increased linearly from root to tip and thus appeared as a segment of a cone. Conical camber originated with Charles Hall in the 6- by 6-foot tunnel, but its study was taken up in other facilities such as the 12-foot tunnel and covered a period of 6 years and tests of more than 25 models.

Among the numerous reports dealing with the conical-camber investigation were RM A55Gl9 by John Boyd, Eugene Migotsky, and Benton Wetzel, and RM A58C21 by Robert Sammonds and Robert Reynolds. To the great satisfaction of Ames engineers, conical camber found immediate and very profitable application to three delta-wing airplanes produced by the Consolidated-Vultee Aircraft Corp. The airplanes thus benefited were the supersonic F-102 and F-106 fighters and the new double-sonic B-58 bomber. The added range and flight endurance which conical camber gave to these airplanes were especially valuable.

Bodies. Although research at Ames was becoming increasingly concerned with the airplane as a whole, a considerable amount of effort was still being spent on important components such as wings and bodies. Bodies had become major lift-producing elements in the configurations of hypersonic aircraft and thus were assuming an increasing importance. The desirability of flattening the fuselage to increase its lifting ability was considered by Leland Jorgensen in TR 1376. Through such a measure, it was felt, the wings might be reduced to stubs or completely eliminated.

Much of the research on bodies that took place at Ames during this period was of a theoretical nature. Contributors included Milton Van Dyke who produced TR 1374, "The Similarity Rules for Second-Order Subsonic and Supersonic Flow," and TN 4281, "Second-Order Slender-Body Theory -Axisymmetric Flow." Another contribution to the theory of flow around bodies was TR 1328, "A Second-Order Shock-Expansion Method Applicable to Bodies of Revolution near Zero Lift," by Clarence Syvertson and David Dennis This report, dealing with hypersonic flows, represented an extension of the generalized shock-expansion method which Eggers, Savin, and Syvertson had presented to the Institute of the Aeronautical Sciences in 1955 and which had appeared in volume 22, No. 4 of the IAS Journal.

In the study of aircraft bodies, interest continued in the problem of [248] defining the body shape that would offer the minimum drag. In order that the body so defined should have some utility, it was necessary to impose certain conditions having to do, for example, with the length, volume, maximum diameter, or base area of the body. Minimum-drag shapes for subsonic and supersonic flight had been fairly well established by 1955, but some uncertainty remained concerning the optimum shape for hypersonic flight. An attempt to solve the hypersonic problem is described in TR 1306 (ref. B-39) by A1 Eggers, Meyer Resnikoff, and Dave Dennis. The approach used by the authors of this report was a combination of theory and experiment. The experimental phase of the program was carried out in the 10- by 14-inch tunnel at Mach numbers ranging up to 6.28.

Area Rule. The flurry of research activity that took place after the discovery of the Area Rule lasted for a number of years and changed in character with time. There was an effort first to confirm the rule experimentally and then to establish a more solid theoretical basis for it. Additionally, there were moves to extend the rule beyond the original transonic speed range and also to determine the limits of its application, particularly as to airplane configuration. These developments led aerodynamicists to a fuller appreciation of the possibilities of manipulating the wave patterns of the various components of an airplane to achieve beneficial results such as drag reduction, lift improvement, or improved inlet efficiency. Instead of enclosing engines, rocket pods, and other components within the structure of an airplane, it now seemed desirable to mount them externally as individual units strategically positioned to produce shock-wave cancellation or lift augmentation. Consolidated's double-sonic B-58 bomber was built on this principle. The optimization of airplane designs using Area Rule methods thus became a fascinating game which, like chess, was a matter of knowing where to place one's pieces.

As mentioned in an earlier section of this volume, R. T. Jones had developed what was known as the Supersonic Area Rule. At the same time the theoretical basis of the rule and the limits and manner of its application were being studied by Max Heaslet, Harvard Lomax, John Spreiter, and Frank Fuller. In TR 1318, Heaslet and Spreiter reviewed three-dimensional transonic-flow theory as related to the Area Rule concept; and, in TR 1282, Lomax and Heaslet provided a special method for applying the Area Rule at supersonic speeds.

Among the experimentalists working on the Area Rule problem were George Holdaway, William Page, John McDevitt, Fred Sutton, and Robert Dickey. The extremes to which the experimentalists went in their studies are revealed in RM A58C03 (ref. B-40), an interesting report authored by George Holdaway and Elaine Hatfield.

Optimization Studies. Area Rule developments aroused the interest of a few aerodynamicists at Ames in the somewhat fanciful question of what would be the minimum possible value of wave drag obtainable by any [249] distribution of the elements of an airplane within a prescribed region. Optimization studies of this kind were reported by Bob Jones in TR 1335, and by Max Heaslet and Frank Fuller in TR 1385 (ref. B-41) . Jones' study went so far as to consider drag-due-to-lift in addition to wave drag.

Optimization studies of a somewhat more immediate value were also underway One line of effort was aimed at the development of an airplane configuration to operate at the highest possible efficiency (highest L/D), at a Mach number of about 3.0. This particular Mach number was one good design step beyond the operational capabilities of current military airplanes and close to the aiming point of a new bomber (B-70) which, late in 1957 was being considered by the Air Force.

The work of Bob Jones and others drew the attention of the Laboratory to arrowhead-wing configurations characterized by long, slender, highly swept spans tapering uniformly to pointed tips. Such wings, properly twisted and cambered and swept behind the Mach cone, appeared to give rather high values of L/D at supersonic speeds. A number of analytical and experimental studies of arrowhead-wing configurations were undertaken during this period. One of these was reported in TN 4361 by Elliott Katzen of the 1- by 3-Foot Tunnel Branch. Elliott found that, with a twisted and cambered arrowhead wing, it was possible to achieve a trimmed (wing in longitudinal balance) lift-drag ratio of as much as 9.0 at a Mach number of 3.0. The addition of fuselage, tail fins, engine pods, etc., of course, could be expected to reduce this value of L/D by several points; nevertheless, at Mach 3.0, such an efficiency was regarded as being very high.

While arrowhead wings had some possible use for airplanes designed to fly at Mach numbers up to 3.0, they were, unfortunately, structurally weak and unsuitable for aircraft such as boost-glide vehicles, which were expected to fly at hypersonic speeds. For hypersonic aircraft, new configurations were needed and also new principles for achieving maximum flight efficiency. The thoughts of Ames engineers on this subject led to the development of the interference method for designing hypersonic aircraft. The method was based on ideas proposed by Vernon Rossow (RM A55L08) and others outside of NACA. However, as reported in RM A55L05 (ref. B 42), it was first seriously applied to airplane design by Al Eggers and Clarence Syvertson.

The interference method as applied at Ames took cognizance of the intense and sharply defined pressure fields cast by the various parts of hypersonic aircraft and was an attempt to so arrange these parts that the interference between the pressure fields they cast would produce certain specific benefits. For example, any portion of the fuselage Iying below the wing would generate a pressure field that would favorably augment wing lift, whereas any part of the fuselage Iying above the wing would have the opposite, unfavorable, effect.

Application of the interference method at Ames resulted in a characteristically flat-topped configuration having no portion of the fuselage above...

 


[
250]

Flat-topped hypersonic airplane configuration. Points at left are dejected wing tips.

Flat-topped hypersonic airplane configuration. Points at left are dejected wing tips.


Clarence A. Syvertson.

Clarence A. Syvertson.

 

....the wing. The fuselage was typically a semicone, uniformly tapered and extending from the nose to the trailing edge of the wing. The wings were swept just within the shock cone of the fuselage and the tips were bent sharply downward to provide some stability and to take greater advantage of the pressure field of the fuselage.

Clarence Syvertson, it should he noted, had been an important contributor to hypersonic-flow research and it was in recognition of his outstanding achievements in this field that in 1957 the Institute of the Aeronautical Sciences presented him with its Lawrence Sperry Award.

Complete Configurations. A number of interesting investigations were made during the mid-1950's of complete configurations representing [251] flyable aircraft. One such investigation, run in the 12-foot tunnel by a team headed by George Edwards, was concerned with the aerodynamic problems involved in the design of a large turboprop-powered, swept wing airplane of the bomber or transport type. Although turbojet engines dominated the aeronautical scene at this stage, there was still a great deal of interest in the use of turbopropeller engines for long-range high-speed transports. But the well-known peculiarities of swept wings combined with the disturbing effects produced by the slipstreams of powerful turboprop tractor engines were expected to generate many unpredictable stability and control problems. Hence it was considered desirable to run tests on a powered model of what Ames engineers considered a practical, perhaps an optimum, configuration for such an airplane. A rather large amount of effort went into this study, which was reported in TN 3789 and TN 3790. Ames engineers were later reassured concerning the validity of their work when the Russians came out with a turboprop transport, or bomber, which closely resembled the configuration that had been found in the 12-foot tunnel to be optimum1.

The movement of the wing toward the rear end of the fuselage appeared to be a trend in the design of supersonic aircraft. This trend arose partly from an increase in the length and fineness ratio of the fuselage and partly from a rearward movement of the center of gravity as the engines were brought closer to the base of the fuselage. In a number of designs the fuselage projected far in front of the wing but only a short distance behind. The question then arose as to whether it would not be better to mount the horizontal tail at the front end of the fuselage rather than, according to tradition, at the rear. At Ames, tail-first arrangements, known as "canard" configurations, were investigated extensively in the 6- by 6-foot tunnel and other facilities. Such studies were said to have led to early conceptions of a supersonic transport airplane. Major contributors in the canard investigations were Victor Peterson, John Boyd, Gene Menees, and Charlie Hall. A summary and analysis of the results of these studies are contained in RM A58D24 (ref. B-43), "Effects of Canards on Airplane Performance and Stability," by Charles F. Hall and John W. Boyd.

One of the important research and development programs undertaken by the Ames Laboratory between 1954 and 1957 concerned boundary-layer ControI (BLC) as actually applied to an F-86 airplane. The program, jointly conducted by the 40- by 80-Foot Tunnel and the Flight Research Sections' covered an investigation of three different kinds of boundary-layer control (1) boundary-layer removal by suction at the leading edge of the wing, (2) boundary-layer removal by suction at the leading edge of the flap, and (3) the use of a jet just ahead of the flap to energize the boundary layer. All forms of BLC were intended to improve the landing characteris-...

 


[
252]

John W. Boyd.

John W. Boyd.

 

...-tics of the airplane and all were intended to exploit the air-pumping capacity of the jet engines which in landing is otherwise largely unused.

In the 40-by-80, where the BLC program was strongly led by Bill Harper, details of the flow system were worked out through the use of a model incorporating a jet engine and the actual wings of the F-86 airplane. The design features thus developed were then applied to an actual F-86 airplane and their performance was checked by flight tests. In the flight tests the landing approach with and without BLC was evaluated by a number of Air Force, Navy, contractor (North American), and NACA pilots.

The BLC program was of rather large magnitude. Its results are summarized in three technical reports. The work on leading-edge suction is reported in TN 1276 by Curt Holzhauser and Richard Bray; and the work on flap suction in TR 1370 by Woodrow Cook, Seth Anderson, and George Cooper. The work on the jet flap, perhaps the most practical arrangement' is reported in TR 1369 (ref. B-44) by Mark Kelly, Seth Anderson, and Robert Innis. Aside from the authors of these reports, there were many contributors to the program, including Jules Dods and Earl Watson who, as reported in RM A56C01, investigated the jet flap in the 7- by 10-foot wind tunnel.

During this period, interest began to develop in airplanes having vertical takeoff and landing (VTOL) and short takeoff and landing (STOL) capabilities. A number of investigations of VTOL or STOL devices were made in the 40- by 80-foot tunnel. These devices included a full-scale helicopter rotor, a wing having a propeller mounted in a circular hole cut through the chord plane, and the McDonnell Model XV-1 Convertiplane which was a combination of helicopter and airplane.

 

[253] PROPELLERS AND INLETS

 

Propeller studies of limited scope continued during this period. In the 40-by-80, Vernon Rogallo and Paul Yaggy completed their study of once-per-revolution stresses that occur in tractor propeller blades operating in the nonuniform flow field of a wing. This study is reported in TR 1295 authored by Rogallo, Yaggy, and John McCloud. Meanwhile work on the performance of supersonic propellers was continuing in the 12-foot tunnels under the supervision of Fred Demele, William Otey, and Carl Kolbe. Representative work conducted by this team, a study of the effect of blunting the trailing edge of a supersonic propeller, is contained in RM A55J12, authored by Demele and Kolbe.

By this time inlet design had developed into a rather sophisticated science. Many people, all over the country, had contributed to it. The problems of inlet design had increased with the speed of airplanes. The main inlet problem arose because the air received by the inlet had to be slowed down 2 to subsonic speeds before it was delivered to the engine. The aim was to accomplish the slowdown with a minimum of external disturbance and a maximum of pressure recovery at the engine face. But it was in the slowing-down process that much of the inefficiency of supersonic inlets arose. A supersonic airstream, unlike a subsonic airstream, does not slow down gradually. It does so in one or more discrete steps in the form of compression shock waves and the final step, often the only step, from supersonic to subsonic speed is always in the form of a shock, called a normal shock that lies directly across the flow. The remaining shock waves, if any, are of the oblique variety.

The pressure rise through a compression shock is diminished by thermodynamic losses which are a function of shock intensity-itself varying in accordance with some moderately high power of the velocity drop across the shock. The pressure diminution, or pressure loss, that results from a given shock-induced slowdown in an airstream is thus less if the slowdown is accomplished through a series of mild shocks rather than through a single intense normal shock. The ordinary scoop inlet, which operates on the basis of a single normal shock, was known to incur pressure losses of several percent at a Mach number of 1.6 and quite intolerable losses at higher Mach numbers. It was clear that, for airplanes of higher speed than M 1.6, some way had to be found to introduce additional, oblique shocks into the inlet airstream.

The oblique shock represents the means by which a supersonic stream changes direction, and it may be produced by enforcing a direction change....

 


[
254]

Schlieren photograph of air inlets showing wedge production of oblique shock waves ahead of inlet at Mach number 1.95.

Schlieren photograph of air inlets showing wedge production of oblique shock waves ahead of inlet at Mach number 1.95.

 

....in the stream. Enforced turning, it was found, could be accomplished externally (ahead of the inlet) by a system of cones or wedges, or internally, as in the diffuser of a supersonic wind tunnel, by a contraction of the internal flow passage of the inlet. The pressure loss of either system was low compared with that of a normal shock inlet, but the internal compression (shock) system had two distinct advantages over the external compression system-and these were of a kind which increased in value with Mach number.

One advantage of the internal compression inlet was that it offered the possibility of eliminating the external drag that unavoidably accompanies an external compression inlet. The second advantage was that, since the direction change it produced was convergent rather than divergent, its overall volume could possibly be made smaller. Unfortunately, the internal compression inlet, like a supersonic wind tunnel, required a variable geometry. This feature, Emmet Mossman found, could he provided by incorporating a translating centerbody. The resulting inlet operated on the basis of a combination of external and internal compression (a combination of divergent and convergent direction changes) and thus was no longer a pure internal compression type. Nevertheless, it constituted an arrangement that was troth practical and efficient.

It was in the investigation of ideas such as those mentioned that a number of men at Ames were engaged for many years. Representative work performed by this group during the current period is reported in RM A56F06 by Emmet Mossman and Frank Pfyl and in RM A55F16 (ref. B-45) by Wallace Davis and Richard Scherrer. The latter report is a comprehensive summary of inlet research conducted at Ames and elsewhere.

 

[255] DYNAMIC STABILITY AND LOADS

 

Since wind-tunnel models required quite a firm mounting to withstand air forces, the conventional wind tunnel had long been regarded as a rather difficult means by which to study dynamic stability. Nevertheless, in such facilities it had proved feasible to allow control surfaces, such as ailerons, a degree of motion so their dynamic performance could be studied. Also, in 1949, the pitch damping characteristics of certain delta-wing configurations had been investigated in the 6- by 6-foot tunnel through the use of springmounted models.

The scope of this work was narrow, however, and, with the development of new airplanes having widely different configurations and operating ranges, the need for the adaptation of wind tunnels for dynamic testing greatly increased. While theory was useful in predicting dynamic stability, a method for obtaining confirming data in wind tunnels was needed. Two groups of men at Ames undertook to solve this problem. Neither group attempted to change the basic wind tunnel-only the model-mounting system.

One group consisting of Henry Lessing, Tom Fryer, and Merrill Mead developed a rather ingenious mounting system in which the model was given two degrees of freedom (roll and yaw) and was forced to oscillate in roll. This device, described in TN 3348, worked fairly well and was able to provide reasonably reliable information on rolling derivatives as well as on certain directional stability and damping in yaw parameters.

The second group working on the dynamic stability problems consisted principally of one man, Benjamin Beam, who was ably supported by the Wind Tunnel Instrument Research Branch. Beam's device, somewhat more sophisticated and elaborate than that of the first group, allowed the model freedom to roll, pitch, or yaw (any two in any one test) as excited by forced oscillations about one of the three axes. The exciting oscillations were controlled by velocity feedback that permitted testing through ranges of variables which, for practical reasons, could not be covered by the more conventional types of oscillatory testing. The processing of data from the device was greatly simplified by the use of analog-computer elements in the strain-gage circuitry. The system was designed primarily for the measurement of damping derivatives, in pitch, roll, and yaw, as well as of cross-derivatives such as rolling moment due to yawing.

Beam's device, which he described in TR 1258 (ref. B-46), was first used in the 12-foot tunnel and later in other tunnels. It provided advance information on the dynamic stability of many new aircraft-even warheads -and thus was extremely valuable. Some of the applications of the device are reported in RM A55A28, RM A56I04, and RM A58F09.

The analog computer was now also being used for dynamic stability Studies. In such use the computer actually simulated, or pictorially reproduced through graphs, the motions of the airplane as affected by control....

 


[
256]

Benjamin H. Beam and dynamic stability spring-mounted model of Convair F-102 airplane.

Benjamin H. Beam and dynamic stability spring-mounted model of Convair F-102 airplane.

 

.....motions, changes in control, stability or inertia parameters, or changes in flight conditions (altitude, etc.). To do this, the computer had first to be programed with certain inputs including (1) the equations of motion; (2) the basic stability derivatives or transfer functions of the airplane, separately obtained; and (3) the calculated or measured changes in the transfer function as produced by those changes in control and stability parameters which were to be investigated. Once the computer was set up, stability and control problems could be studied easily, cheaply, and with no risk of life or property. An example of the use of the analog computer for dynamic stability studies is given in RM A56H30 by Brent Creer. The nature of the study is indicated by the title, "An Analog Computer Study of Several Stability Augmentation Schemes Designed to Alleviate Roll-Induced Instability." The characteristics of the North American F-100 fighter were used in this investigation.

During this period, as always, a number of studies were undertaken at Ames on the subject of airloads. In RM A55C02, Perry Polentz, William Page, and Lionel Levy of the 2- by 2-foot tunnel reported on a wind-tunnel investigation at Mach numbers up to 0.9 of the unsteady normal-force characteristics of a series of 27 representative airfoil sections. Also, as reported in TN 3346 and TN 3500, John DeYoung, aided by Walter Barling, continued the span-load studies which he had begun years before. In the 6-by 6-foot tunnel, Dave Reese undertook an investigation, reported in RM A55F01, of the unsteady lift induced on a wing in the downwash field of an oscillating canard control surface. This was another instance of the kind just mentioned in which problems of dynamic stability or airloads were studied by forced, or free, oscillations of a control surface.

Murray Tobak had by this time become quite well known for his theoretical studies on the dynamics of airplanes and missiles. One of his contributions during this period was a study, reported in TN 3290, on the [257] minimization of airplane response to random gusts. In addition to giving mathematical descriptions of the responses of airplanes to gusts, Murray in this paper considered the problem of minimizing the response and derived the theoretical requirements of a compensating force system.

 

FLIGHT RESEARCH

 

Simulators. A new trend had begun in flight research. This movement was toward the use of simulators for studying aircraft guidance and control problems. The simulators were of two kinds: airborne and ground based. Work on airborne simulators had actually begun several years earlier with the development of the variable-stability airplane. The first variable-stability airplane was the propeller-driven F6F. The converted F6F had proved very useful and a more modern airplane, the F-86 swept-wing jet fighter, had then been converted for variable-stability work. It, also, proved very useful. Ames' work with variable-stability airplanes was indeed responsible for the use of negative dihedral in Lockheed's sleek new doublesonic F-104 fighter. Negative dihedral had perhaps never before been used in an airplane, and the concept of its use was quite foreign to the thinking of airplane designers. It was therefore with considerable astonishment that Kelly Johnson, Lockheed's chief designer, heard that an F-104 simulation at Ames had shown negative dihedral to be distinctly advantageous. He immediately dispatched Tony Lee Vier, Lockheed's crack test pilot, to the Laboratory. Tony flew the variable-stability airplane and confirmed the Ames finding. The upshot was that negative dihedral became one of the distinctive features of Lockheed's outstanding F-104 airplane.

Other uses of the variable-stability airplane at Ames during this period are reported by Walter McNeill and Brent Creer in RM A56C08, "A Summary of Results Obtained During Flight Simulation of Several Aircraft Prototypes With Variable Stability Airplanes." Creer had been project engineer in the conversion of the F-86 airplane with which much of the reported test work was done.

Another flight simulator developed and used at Ames consisted of an F-86 airplane in which a variable-control system had been installed. This device allowed the test pilot to vary the dynamic characteristics of the airplane's longitudinal control system over a wide range and thus to study such factors as control feel, sensitivity, breakout force, and time constant. Research applications of the variable-control-system airplane are reported in RM A57L10 by Norman McFadden, Frank Pauli, and Don Heinle.

The research story for the previous period (1950-1953) described the tracking investigation made by Ames flight engineers with several airplanes equipped with cameras and fixed gunsights. The next phase of this program was to perform tracking tests with an F-86 airplane equipped with a gyromounted, electronically controlled, computing gunsight that would automati-[258] -cally provide the proper "lead" on the target as dictated by its course, speed, and distance. The accuracy with which a pilot could track a maneuvering target airplane with this additional automatic element in the control system was thus determined and reported by Burnett Gadeberg and George Rathert in RM A54K16.

A target airplane had always been used in the tracking tests at Ames, but this practice created added trouble and expense and there was also some danger of collision between the two airplanes. Ames engineers had become so adept in their automatic-control-systems work that they felt it would be possible to simulate the maneuvering target airplane with a special arrangement of equipment that could be carried in the tracking airplane. This system was duly developed and described by Brian Doolin, Allan Smith, and Fred Drinkwater (pilot) in RM A55F20, "An Air-Borne Target Simulator for Use in Optical-Sight Tracking Studies." The target airplane in this simulation appeared as a moving spot of light on a glass screen in front of the pilot's windshield. The system was very complex but it worked well except that the spot of light which simulated the target airplane did not give the forewarning of a turn that was provided by the banked wings of an actual target airplane.

The target simulator developed by Ames engineers was shortly adapted for simulations of the guidance of the radio-controlled Bull Pup missile that the Navy had developed. This missile was designed to be air-launched and, through a radio-control link, visually guided by the pilot to a ground target. Since the Bull Pup guidance was a tracking job of sorts, it could be simulated quite readily through a modification of the Ames tracking simulator. The results were highly successful and the resulting simulator proved very useful in training Navy pilots in Bull Pup missile operations. The Bull Pup simulator and its use were described by Joe Douvillier, John Foster, and Fred Drinkwater in RM A56G24 (ref. B-47), "An Airborne Simulator Investigation of the Accuracy of an Optical Track Command Missile Guidance System." Along the same line, Bill Kauffman, who had contributed much to the simulator work at Ames, prepared a paper on "Flight Applications of Target Simulator Principles" which was published in the November 1955 issue of the IAS Aeronautical Engineering Review. For his flight simulator work, Kauffman, in 1955, received from the Washington, D.C., Junior Chamber of Commerce the Chamber's Arthur S. Flemming Award

The ground-based simulator studies at Ames began with a short investigation of the longitudinal stability and control characteristics to be expected of the X-15 airplane as it traveled through the upper reaches of-essentially out of-the sensible atmosphere. In this study, which was carried out by Howard Matthews and Robert Merrick and reported in RM A56F07, a pilot was incorporated in a closed-loop simulator which also included a simple control panel, a control stick, and an analog computer programed to represent the dynamics of the airplane. The control panel and control stick [259] were mounted on a table in front of which the pilot sat. Thus the cockpit representation was rather crude. Nevertheless, some useful information, including pilot's opinions, was obtained regarding the pilot's ability to control the airplane as it departed from and returned to the earth's atmosphere.

Perhaps the next use of a ground-based simulator at Ames was in a study, reported by Maurice White and Fred Drinkwater in RM A57D30, of the approach speeds of carrier-based aircraft. The objective of the study was to find out exactly what it was in the stability and control characteristics of an airplane that caused a pilot to select a certain approach speed in a carrier landing. Here again pilot opinion was a dominant element of the study and the opinions of numerous pilots were elicited. The simulator was again of the simplest kind, similar to the one used in the X-15 study, with a patchwork of gadgetry and the pilot sitting on a stool "flying" an airplane onto a carrier deck. The airplane and carrier deck were represented in simulated form on an oscilloscope screen. The simulations covered four of the Navy's new jet fighters and the results obtained seemed to confirm the value of ground-based flight simulators as a research tool. The validity of the simulator results was in this case confirmed by a flight program in which the carrier-approach handling characteristics of 41 different types of fighter airplanes were obtained together with extensive pilot opinion data. The flight program was reported by Maurice White and Bernard Schlaff in RM A57L11.

At this stage, ground-based flight simulators failed by a wide margin to completely simulate the flight of an airplane and it seemed too much to expect that complete simulations would be possible. It was clearly desirable to simulate the important elements for each study-if one only knew what they were. One obviously incorrect feature of the simulations was that the pilot was sitting still and thus his voluntary and involuntary responses to the motions and attitudes of the airplane in flight were not excited. This omission perhaps had little effect in the carrier-approach studies where the simulated airplane was presumably in a long steady glide. But what if the problem under study was that of pitchup, in which the motions of the airplane not only provide the pilot with a cue for corrective action but throw him around a bit? For such studies the pitching motion should certainly be represented-but how? This was a problem of interest to Melvin Sadoff, who had done so much work on the pitchup problem.

Mel found that unfortunately there was no simple inexpensive way to incorporate even one degree of motion into a ground-based flight simulator. He discovered, nevertheless, that it was possible to use a Link trainer to obtain certain empirical data on pilot response to pitching motion which could later be used in analog studies of the pitchup phenomenon. This procedure was followed in a study reported in RM A55D06, "A Method for Evaluating the Loads and Controllability Aspects of the Pitch-Up Problem," by Melvin Sadoff, Frederick H. Matteson, and C. Dewey Havill. Shortly after this investigation was completed, there was undertaken at

[260] Ames a design study of a pitch-roll chair-a motion generator that could be used in future simulator studies. This then was the beginning of the Laboratory's work on ground-based, piloted flight simulators.

Pilot Opinion. The interpretation of pilot opinion was clearly becoming increasingly important in flight research conducted either with airplanes or with ground-based flight simulators. For a number of years George Cooper had given considerable attention to this matter. To him, the problem of obtaining consistent and useful data from a group of pilots had four important elements. First was the formulation of the question-the choice of the particular factor upon which pilot-opinion data were to be accumulated and the elimination of all ambiguity with regard to the question. Second was answering the question in words the meaning of which had been agreed upon. A standardized rating system was required. Third was weighing the answers of the pilot, taking into consideration his experience, adaptability, and current duties. Fourth was the use of ground-based piloted flight simulators for comparison with flight so that the pilot could conveniently study his own reactions in safety and could also evaluate the importance of what had been left out of the simulation. A much more complete description of George's thoughts in these matters is contained in his paper "Understanding and Interpreting Pilot Opinion" (ref. B-48) which in March 1957 was published in the IAS Engineering Review. A notable contribution contained in this paper was Cooper's Pilot Opinion Rating Scale, which became widely accepted as a standard in flight-research circles.

George Cooper, a quiet competent individual and an excellent engineering test pilot, was much liked and highly respected by his colleagues. They felt that Cooper, because of his unassuming character, had never received the credit he deserved and pointed out that back in 1948-1949, while the heroic, much lionized pilots of the X-l were occasionally breaking the sound barrier at Muroc, he and Rudy Van Dyke at Ames were doing it twice a day in an F-86 without any publicity whatsoever. George's fine work received official recognition in 1954, when he won both the Octave Chanute and the Arthur S. Flemming Awards.

Automatic Systems. Work on automatic control systems for aircraft continued at a good pace during this period. A fairly large part of this effort had to do with the optimization of systems used for guidance of missiles and automatically controlled interceptors. The missile-system studies can perhaps be represented by two investigations: one was covered in a paper by Marvin Shinbrot entitled "Optimization of Time-Varying Linear Systems with Non Stationary Inputs," published in the ASME Journal, volume 80, No. 2; and the other was presented in RM A55E11, "Application of Statistical Theory to Beam-Rider Guidance in the Presence of Noise. I-Wiener Filter Theory," by Elwood Stewart. This paper was an important first application of Wiener's filter theory to the elimination of certain spurious radar [261] signals, called "noise" or "glint," which often caused serious, if not completely defeating, errors in missile guidance systems.

Ames research on automatic interceptor systems during this period is well represented by two studies. One of these was carried out by Stanley Schmidt and William Triplett and reported in TN 3387, "Use of Non-Linearities to Compensate the Effects of a Rate-Limited Servo on the Response of an Automatically Controlled Aircraft." The other study was reported by Bill Triplett and Francis Hom in RM A57D09, "Flight Tests of an Automatic Interceptor System with a Tracking Radar Modified To Minimize the Interaction between Antenna and Interceptor Motions." It can be said that Bill Triplett's contributions in this area of research at Ames had become very large and extremely important.

Dynamic Response. Investigations of the dynamic response of airplanes were also carried on during this period. Of particular note was a study undertaken jointly by Ames and the High Speed Flight Station at Edwards. This study, carried out by Henry Cole, Stuart Brown, and Euclid Holleman (of HSFS), was reported in TR 1330 (ref. B-49), "Experimental and Predicted Longitudinal and Lateral-Directional Response Characteristics of a Large Flexible 35° Swept-Wing Airplane at an Altitude of 35,000 Feet." The airplane in this case was a B-47 which had been very thoroughly instrumented to measure the dynamic response of various parts of the airplane structure to excitation in the form of pulsed control motions. The measured response motions of the aircraft were converted to frequency response and transfer functions, which were compared with analytically predicted values. The study was of particular interest because it gave an insight into the dynamic problems of the highly flexible bomber and transport airplanes then being built.

 

BOUNDARY LAYER, SKIN FRICTION, AND AERODYNAMIC HEATING

 

The boundary layer, that thin reaction zone between an airplane and its external environment, had always been of interest to aerodynamicists and airplane designers. At first the interest lay mainly in its potentialities for causing flow separation with accompanying loss of lift and high-pressure drag. Soon it had assumed additional importance as a result of its effect on skin friction. More recently, the boundary layer had acquired an extremely important new meaning relating to its effect on aerodynamic heating; and now great significance was being attached to its chemistry and thermochemical reactions with the aircraft. Inasmuch as the significance of none of the earlier interests had lessened in any large degree, the overall importance of the boundary layer had now reached impressive proportions. It is not surprising therefore that, during this period, the boundary layer and its related influences were important items for study at the Ames Laboratory.

 


[
262]

Shadowgraph of boundary-layer transition at Mach 3.

Shadowgraph of boundary-layer transition at Mach 3.

 

In an interesting investigation made in the supersonic free-flight tunnel, Carlton James, as reported in TN 42.35, demonstrated again that the boundary layer did not change suddenly from laminar to turbulent, but rather did so through an intermediate series of bursts of turbulent flow. The same boundary-layer transition characteristic was found in the 12-foot tunnel where, as mentioned in RM A56G17, Fred Boltz and his partners were able to detect transition by means of a microphone. Another significant study of boundary-layer transition, on a body of revolution, was made in the supersonic free-flight tunnel by James Jedlicka, Max Wilkins, and Al Seiff (TN 3342).

Meanwhile, Dean Chapman was continuing studies, begun years earlier, of separated flows. In TR 1356 (ref. B-50), Dean, together with Donald Kuehn and Howard Larson, presented the results of an investigation of the effect of the location of the transition point on the character of separated flows. Chapman also, as described in TN 3792, made a theoretical analysis of heat transfer in regions of separated flow.

As much of the heating of an airplane's surface was known to be caused by the viscous frictional forces of the boundary layer, considerable effort was spent at Ames on the difficult technique of actually measuring the skin-friction force on an isolated surface element by means of a cleverly designed, extremely delicate balance. This notable work was performed by Donald Smith and John Walker of the 12-foot tunnel and reported in TN 4231 (ref. B-51) .

The problem of predicting the rate of heat transfer from the boundary layer to a body as affected by Mach number and body-wall temperature was made much easier by an extremely useful report, TN 4236 (ref. B-52), turned out by Thorval Tendeland late in this period. Earlier, a team comprising Glen Goodwin, Marc Creager, and Ernest Winkler had, in the old [263] 6-inch heat-transfer tunnel, investigated the local heat-transfer and pressure-drag characteristics of a yawed circular cylinder, representing the blunt leading edge of a swept wing. It was well known then that the heat transfer to sharp leading edges was so intense that the edges would quickly melt off -better, it was felt, to make them blunt to start with. Sweep could then be used to reduce the drag penalty caused by blunting as well as to further reduce aerodynamic heating. This investigation, by Goodwin, Creager, and Winkler, was reported in RM A55H31 and is said by Ames engineers to have significantly influenced the design of the X-15 hypersonic research airplane. In an extension of the study, reported in TN 4142, Marc Creager investigated the heat-transfer and pressure distributions over flat plates (representing wings) with blunt, swept-back leading edges. This TN, one of a series of reports by Creager on similar configurations, was the first to issue from the new 8-inch low-density tunnel and contains a description of that facility.

Although the heating of the wings of airplanes represented a very difficult problem, it was less critical than the heating of the blunt noses of ballistic missiles or of spacecraft returning from extra-atmospheric flight. The problem of the ballistic missile was all the more difficult because the flow pattern about the blunt nose of such a vehicle was extremely complex and did not lend itself to theoretical treatment. Indeed, the "blunt-body problem" was one over which theoretical aerodynamicists throughout the country lost considerable sleep. They could thank Harvey Allen for that. Much credit, then, is due to a team composed of Al Eggers, Fred Hansen, and Bernard Cunningham, who were able to develop an approximate theory for predicting the rate of heat transfer to the stagnation region of blunt bodies. This analysis was reported in TN 4229.

The supersonic free-flight tunnel had proved to be a very fine instrument for investigating the boundary layer, skin friction, heating, and stability of missiles, but achievable Mach numbers were still too low to simulate ballistic-missile conditions fully. Nevertheless, a large number of studies relating to ballistic missiles were made in the facility. Among these investigations was a study of the boundary-layer-transition characteristics of high-drag bodies of revolution, which was reported in RM A56I05 by Alvin Seiff, Simon Sommer, and Thomas Canning. There was also a very important study by Si Sommer and Barbara Short of turbulent skin friction in the presence of severe aerodynamic heating at Mach numbers up to 7.0. This study was described in TN 3391 (ref. B-53) .

Another very significant study in the SSFF tunnel, reported in RM A57C25 (ref. B-54) by Tom Canning and Si Sommer, was an investigation of boundary layer transition on flat-faced bodies of revolution at Mach numbers up to 9.0. The results related directly to the design of blunt-nosed ballistic missiles and spacecraft. Allen and Eggers had shown that the heating of ballistic-missile warheads could be reduced by making them blunt, thus [264] increasing their pressure drag and bow-shock-wave strength. It was also clear that the heating could be reduced by lowering the skin friction, and one important way of doing this was to induce the boundary layer to remain laminar. The nose of the warhead obviously had to be blunt, but was it possible that the maintenance of a laminar boundary layer would be favored by one form of bluntness over some other? The tests reported by Seiff et al. in RM A56I05 had not been very encouraging with regard to the expectation of obtaining much laminar flow over some forms (round-nosed cones) of blunt body. But perhaps, if the blunt nose was absolutely flat, like the top of a fencepost, or nearly flat, laminar flow might be achieved, and heating reduced, over the whole flat face. This was what Sommer and Canning actually found and reported in RM A57C25.

The prototype version of the atmosphere-entry simulator was put into operation early in 1957; the first research investigation made in it involved tests of copper-clad flat-faced cylindrical test models that were launched by means of the single-stage shock-compression gun at speeds of 14,000 feet per second. The tests were of limited scope; however, spectroscopic measurements of the glowing wake were made, and the models were recovered and examined for erosion or loss of material from the front face. The results, though not overly significant, did nevertheless have some bearing on the design of IRBM nose cones. They also confirmed the feasibility of the AES test method. The study was carried out by Stanford Neice, James Carson, and Bernard Cunningham and was reported in RM A57I26.

Earlier findings had suggested that the first step to be taken in minimizing the aerodynamic heating of a hypervelocity vehicle was to choose the right nose shape. The use of the best temperature-resistant high-strength materials was likewise obvious. Some heat could also be dissipated from the body by radiation to outer space but, since radiation varies as the fourth power of the temperature, it appeared that the body would not become a very effective radiator until surface temperatures rose beyond the endurance of most structural materials. Thus as speeds and heating increased, resort had to be taken to artificial cooling methods. It was recognized that a cooling fluid could be circulated within the body or extruded into the boundary layer through pores or small openings in the surface of the body. The latter method was called transpiration, or sweat, cooling and, since it dissipated material that had to be carried in the vehicle, it was also known as ''mass transfer" cooling.

Transpiration cooling, it was realized, could be very effective. First of all, the cool transpired fluid carried off some of the body heat. Second, it tended to insulate the body from the hot boundary layer; and third, it reduced skin friction. There were, moreover, other beneficial effects of a more subtle nature. The magnitude of these benefits depended greatly on the nature of the transpiration fluid. It could be air, of course, and also water. Water was effective because it absorbed heat in changing to a gas. Aside [265] from these two choices, the transpiration fluid could be one of the lighter gases, such as helium or hydrogen, which had certain properties that were suitable for the purpose.

At Ames, transpiration cooling became the subject of considerable attention A typical report resulting from this work is RM A56D05 by Thorval Tendeland and Arthur Okuno. This report provides experimental evidence of the effect of cool-air injection on the skin friction of a turbulent boundary layer. Another example is TN 4149 by Morris Rubesin and Constantine Pappas, which contains a theoretical analysis of the effect on skin friction and heat transfer of injecting either helium or hydrogen into a turbulent boundary layer. The latter report shows that the injection of helium and hydrogen, particularly of hydrogen, is several times as effective in reducing both skin friction and heat transfer as the injection of air. It thus indicated the advantage of using a cooling fluid of low molecular weight. Of course the use of hydrogen is perhaps only a theoretical concept as it would burn if used to cool the external surfaces of an air vehicle.

As the air in the boundary layer and behind the bow wave of a hypervelocity body becomes heated, its thermodynamic and chemical properties change. Properties such as thermal conductivity, heat capacity, and viscosity vary with temperature and, as the temperature continues to rise, the two atoms that make up a molecule of oxygen or nitrogen become so agitated that they fly apart in a heat-absorbing chemical process called "dissociation." Later, when the air begins to cool, they recombine, a process that generates heat. If the temperature of the air rises much above the dissociation level, the atoms themselves attain such violent motions that some of their electrons are torn loose and float freely within the body of gas; in this state, the gas is called a "plasma." Inasmuch as the negatively charged electrons and the positively charged nucleus of an atom are normally in electrical balance, the removal of an electron leaves the remainder of the atom with a positive charge. Thus a plasma, containing many separate positively and negatively charged particles, is electrically conducting and subject to the influence of magnetic fields.

It is understandable that the aerodynamicists and physicists at Ames were now becoming very much interested in the properties of air at high temperature and that they should seek to expand the knowledge on this subject which was of limited extent, inaccurate, or nonexistent. One of the major contributions in this field was made by Fred Hansen and reported in TN 4150 (ref. B-55) . In this report, Fred computed and tabulated eight or more useful properties of air for temperatures ranging up to 15,000° Kelvin. A somewhat more modest study on "Effects of Oxygen Recombination on One-Dimensional Flow at High Mach Numbers" was made by Steve P. Heims.

Another very interesting study, in a somewhat related field, was made by Vernon Rossow (ref. B-56). Rossow examined theoretically the possi-...

 


[
266]

C. Frederick Hansen.

C. Frederick Hansen.

 

....-bility of reducing skin friction and heat transfer in a flow of plasma over a flat plate by applying a transverse magnetic field. He concluded that the skin-friction and heat-transfer rate could be reduced by this method; whether or not the attempt to do so would be practicable was another matter.

 

BALLISTIC AND BOOST-GLIDE VEHICLES

 

A fairly serious question had arisen at Ames regarding the aerodynamic stability of the blunt-nosed configurations which Allen and Eggers had recommended for the warheads of ballistic missiles. It was necessary for the blunt and otherwise heat-protected face of the warhead to remain pointed forward. Any wild oscillations or tumbling of the reentry body would affect the aerodynamic heating and possibly also its trajectory and its internal workings. Then too, the idea of a man-carrying ballistic missile had appeared and human physical endurance placed another limit on the oscillations and tumbling of a reentry body.

The subject of warhead stability was investigated both experimentally and analytically. In the SSFF tunnel, as reported in RM SA57K18, the use of a single-stage shock-compression gun made it possible to run dynamic-stability tests at Mach numbers up to 15. The study covered by this report was made by Simon Sommer and demonstrated the static stability of one specific warhead design.

One of the early analytical approaches to the reentry stability problem was reported by Harvey Allen in TN 4048, "Motions of a Ballistic MissiIe Angularly Misaligned With the Flight Path Upon Entering the Atmosphere and Its Effect on Aerodynamic Heating, Aerodynamic Loads and Miss Dis-[267] tance." A somewhat more thorough analysis, jointly undertaken by Murray Tobak and Harvey Allen, was reported in TN 4275, "Dynamic Stability of Vehicles Traversing Ascending or Descending Paths Through the Atmosphere." These reports took into consideration the peculiar pseudo-stability that a reentry vehicle acquires by virtue of the fact that the air density along its path is continuously increasing. For an ascending vehicle, a pseudoinstability was evident.

At this time much interest had risen at Ames and elsewhere concerning the usefulness of rocket-launched man-carrying gliders for long-range hypersonic flight in the outer reaches of the atmosphere. NACA had undertaken discussions with the Air Force on the desirability of building an experimental boost-glide vehicle as the next step beyond the X-15 in the research-airplane program. The Air Force was much interested because of the military potentialities of such an aircraft; and, as 1957 ended, it was about to initiate Project Dynasoar which called for the development of a world-girdling, man-carrying boost-guide vehicle.

Ames was already engaged in configuration studies for aircraft of the boost-glide type. Notable examples of these studies are to be found in RM A55E26, "Some Aspects of the Design of Hypersonic Boost-Glide Aircraft," by Alvin Seiff and H. Julian Allen, and the earlier mentioned RM A55L05, "Aircraft Configurations Developing High Lift-Drag Ratios at High Supersonic Speeds," by A. J. Eggers and Clarence A. Syvertson. Eggers and Syvertson proposed some highly swept configurations the design of which was based on elemental considerations of pressure-field interference and aerodynamic heating.

For transporting weapons and people over long distances at high speeds, there now appeared to be several possibilities: a ballistic missile with nonlifting reentry body, a boost-glide vehicle, and, of course, the vastly slower, but possibly more efficient, supersonic airplane. What was really needed was a comparative systems study to evaluate the feasible range of application as well as the principal design and operating problems for each type of vehicle. Such a study, at least as far as the basic elements of the subject were concerned, was undertaken by Alfred Eggers, Harvey Allen, and Stanford Neice and reported in TR 1382 (ref. B-57).

The study made by Eggers, Allen, and Neice, carried out during 1955 and 1956, was very timely and created much interest in aeronautical circles throughout the country. It revealed a number of interesting facts. On an efficiency basis, the study showed, the ballistic and boost-glide vehicles could Compete quite favorably with supersonic airplanes at ranges equal to or greater than half the circumference of the earth. The study also showed that while a ballistic vehicle, sufficiently blunt, could protect precious cargo from the rigors of aerodynamic heating, such a vehicle would, nevertheless, have some limitations as a human carrier owing to excessive reentry deceleration. These limitations, however, which were shown to depend on reentry angle, [268] would not apply at ranges equal to or greater than half the circumference of the earth. Nor would they apply at very short ranges.

The boost-glide vehicle, it appeared, would have a number of advantages over the ballistic vehicle. Its lifting ability and high lift-drag ratio (L/D) would allow it to achieve a greater range than the ballistic vehicle for a given initial boost velocity. Also the boost-glide vehicle could control its flight-path angle and thus its heating rate; its maneuverability in landing was regarded as an obvious additional advantage. It was found, however, that the benefit of lift in increasing the range of the boost-glide vehicle would tend to disappear at ranges approximating the circumference of the earth. At such ranges, owing to the near-satellite speeds involved, lift would lose significance as its role was taken over by centrifugal force. Clearly, with just a little more speed, the vehicle would be in orbit and its range would then no longer be dependent on lift or, for that matter, on drag.

The boost-glide vehicle, it was pointed out, could remain in the atmosphere or be boosted beyond the atmosphere later to return. For flight within the atmosphere, low drag, high L/D, and minimal aerodynamic heating were obviously desirable conditions. These conditions, it was felt, could best be achieved with a wing having blunt, highly swept leading edges. Of course, even with a blunt and swept leading edge, the wing would get very hot; but the heating could be minimized by controlling the rate of descent, and the same method could be used to provide time for a large part of the vehicle's heat burden to be dissipated by radiation into the surrounding sky.

Seiff and Allen, in RM A55E26, looked into the radiation matter and calculated the wing equilibrium temperatures that would obtain when the convective heat input and the radiative heat output were in balance. They found that, for a blunt-nosed, highly swept wing, at a Mach number of 7 and an altitude of 120,000 feet, the equilibrium temperature of the leading edge would be about 2000° F. The remainder of the wing would be cooler. The upper surface would have an average temperature of about 700° F. and the lower surface 1400° F if the flow were laminar and 1900° F if it were turbulent. The higher temperatures found on the lower surface resulted largely, of course, from increased friction arising from the higher air pressure applied to that surface as part of the lifting process.

The importance of Allen's blunt-nose principle was now being recognized and his other major contributions to the development of hypervelocity vehicles were also widely appreciated. In 1955 he received the Sylvanus Albert Reed Award of the Institute of the Aeronautical Sciences and in 1957 he was awarded NACA's Distinguished Service Medal. Allen was also honored by being invited to present the prestigious Wright Brothers Lecture in Washington, D.C., on December 17, 1957, the 54th anniversary of the Wright Brothers' first flight. The subject of the paper (ref. B-58) that he presented on that occasion was "Hypersonic Flight and the Re-Entry Problem."

 


[
269]

Takeoff of an early model Atlas.

Takeoff of an early model Atlas.

 

[270] It was clear that the work of the Laboratory, in terms of speed and altitude, was reaching out to the very edge of space. Indeed, by the end of 1957, rather serious consideration was being given to the problems of space flight. One of the first studies of this kind to be carried out at Ames was made by Fred Hansen in 1956 and reported in TN 3962 (ref. B-59). Fred was concerned with the heating and resulting erosion of the surfaces of space vehicles as they reentered the earth's atmosphere at speeds up to the escape velocity of 37,000 feet per second. Facilities in which tests at such speeds could be made were unavailable, but Fred recognized that in nature such tests were, in effect, being made continuously. He felt that, through the study of meteor flight and meteor trails, much could be learned that would apply to the problem of returning spacecraft. Fred's study was of limited scope, but it represented an interesting concept and produced some useful results.

Some time later, another subject relating to meteors and spacecraft became of interest. A question had arisen as to the extent and nature of the damage to the skin and structure of spacecraft when struck by the small meteoroids which in great numbers whiz through space. Although usually tiny, these meteoroids, because of their tremendous speed (from 7 to 50 miles per second), might well do a lot of damage. There was in fact very little information available on the character of the crater that would be formed from the impact of an object traveling at such speeds. However, the light-gas guns and ballistic ranges developed at Ames offered a means for an introductory investigation of this interesting physical phenomenon. Bullets of any desired shape or composition could be fired into targets of any selected material at speeds up to about 20,000 feet per second. Such test speeds were far from meteor speeds but were the best that could be achieved at the moment. Such a study was undertaken by Alex Charters and G. S. Locke and reported in RM A58B26. In this investigation, which was of limited scope, small spheres made of a variety of metals were fired into targets composed of either copper or lead. Some effort was made to correlate the observed cratering with impact theory. Obviously the whole subject had only been scratched.

Although many-perhaps half-of the Laboratory's recent projects had in some degree been related to space flight, the two just mentioned were among the few that dealt exclusively with the subject. The pace of space research was quickening, however, and it acquired additional impetus from the launchings of Sputniks I and II.

 

* * *

 

One might at this time reflect with some amazement on the rapid increase in man's speed of travel. In the 42 years of aeronautical history prior to the end of World War II, man had achieved a speed of travel of about 650 miles per hour. In the 12 years following World War II, his pace had tripled and at this increasing rate could be expected to grow fivefold or more in the next five years. Truly we were living in a new world of speed.

 


1 Tupolev TU-20, otherwise known as "Bear."

2 Aeronautical engineers, through an odd mental quirk and perhaps through the conditioning of wind-tunnel experience, always think of the airplane as standing still with the air flowing pass it.


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