Quest for Performance: The Evolution of Modern Aircraft
Chapter 5: Design Refinement, 1939-45
Aerodynamic Problems and Refinements
[103] A vast amount of aerodynamic research was conducted in the United States, Great Britain, Germany, and Italy during the years of [104] World War II. No attempt will be made to give a complete summary or abstract of this work; however, a few examples taken from NACA research may serve to indicate the flavor of the activity. More detailed accounts of the research in aerodynamics may be found in references 49, 56, and 104.
Airfoils and High-Lift Devices
The low drag coefficients achieved by internally braced monoplanes equipped with retractable landing gears suggested that any further large reductions in drag could only be achieved through the maintenance of extensive laminar flow over the surfaces of the aircraft. The boundary-layer flow of contemporary aircraft was essentially all turbulent; and since the skin friction coefficients for turbulent flow are much higher than those for laminar flow, the achievement of laminar flow on the surface of the aircraft would be expected to yield large reductions in drag. For example, the skin friction coefficient on a flat plate is reduced by a factor of almost 2 as the point of transition from laminar to turbulent flow is moved from the leading edge to the 50-percent-chord location. In the late 1930's, NACA's Langley Memorial Aeronautical Laboratory undertook the development of special airfoils designed to achieve extensive regions of laminar flow. The problem involved extensive theoretical and experimental investigations and the development of an entirely new low-turbulence wind tunnel. The early work on laminar-flow airfoils is described by Jacobs in reference 76, which was originally published in June 1939. The development of laminar-flow airfoils continued throughout the years of World War II and for several years thereafter. Over 100 different airfoils were derived. The characteristics of these airfoils were published in summary form in reference 18, and a complete exposition of airfoil theory and presentation of airfoil aerodynamic characteristics are given in reference 17.
The profile shapes of two NACA low-drag airfoil sections compared with a conventional airfoil are shown in figure 5.1. The airfoils designated as NACA 661-212 and NACA 631-412 are the laminar-flow, or low-drag, sections; and the airfoil designated as NACA 23012 is a conventional airfoil designed during the 1930's. The 661-212 airfoil was designed to maintain laminar flow to the 60-percen t- chord point, and the 631-412 was designed to maintain laminar flow to the 30-percent-chord point. The designation system used for these airfoils, as well as older conventional NACA airfoil sections, is described in reference 17. As compared with the conventional section, the laminar-flow sections are seen to have the point of maximum thickness located...

chart comparison of 3 airfoil sections
[105] Figure 5. 1 - Shapes of two NACA low-drag airfoil sections compared with NACA 23012 airfoil section.

...farther aft along the chord of the airfoil. The aft location of the maximum thickness point is associated with the need to achieve a particular type of airfoil-surface pressure distribution and is also desirable from the point of view of structural design. A comparison of the section drag characteristics of the NACA 631-412 airfoil and the NACA 23012 airfoil is shown in figure 5.2, in which the drag coefficient is plotted as a function of the lift coefficient for the two airfoils in both the smooth and rough condition.
The bucket in the drag curve for the 631-412 airfoil corresponds to the lift coefficient range in which laminar flow is achieved. In the rough condition, the drag characteristics of the conventional and laminar-flow airfoils are very similar. The roughness employed in the test was a sandlike material that was intended to fix transition near the leading edge in a manner corresponding to a rough and poorly maintained airplane wing. The North American XP-51, which flew in prototype form in 1940, was the first aircraft to employ a laminar-flow- type airfoil section, and most subsequent high-performance aircraft designs utilized these airfoils. One of the essential requirements for achieving laminar flow is that the surface of the wings be manufactured and maintained in an extremely smooth and fair condition. (The term "fair" means that the wing surfaces must be essentially free from waves, that is, ripples, and must conform very closely to the specified...

graph comparison of airfoil sections
[106] Figure 5.2 - Drag characteristics of NACA low-drag and conventional airfoil sections with both smooth and rough leading edges.

...contour shape.) This requirement could be met with highly accurate wind-tunnel models. Unfortunately, methods of aircraft manufacture and maintenance during World War II, and even today, were such that only very small regions of laminar flow located near the leading edge of the wing could be achieved on practical operational aircraft. As a consequence, the use of NACA laminar-flow airfoil sections has never resulted in any significant reduction in the drag as a result of the achievement of laminar flow. A practical means for achieving extensive regions of laminar flow under everyday operating conditions remains a problem today and is still one of the great unsolved challenges in aeronautical research. The NACA low-drag airfoils have seen extensive use and continue to be used on high-performance aircraft because they have better characteristics at high subsonic Mach numbers than conventional airfoil sections. The effectiveness of the NACA laminar-flow airfoils as a means for delaying the adverse effects of compressibility at high subsonic Mach numbers is a classic example of a new technical [107] concept developed to solve one problem but proving highly useful in the practical solution of an entirely different one. Figures 5.1, 5.2, and 5.3 were taken from the unpublished proceedings of a NACA conference held in September 1946 for the purpose of providing representatives of the general aviation industry with the results of previously classified technical data generated during the World War II years.
As the wing loadings of high-performance military aircraft steadily increased, the desirability of maintaining the stalling speed within acceptable limits dictated the need for extensive work on high-lift devices to increase the maximum lift coefficients of aircraft. The types of trailing-edge flaps used in the mid- to late 1930's were usually of the simple plain or split type. For example, the Douglas DC-3 employed simple split-type flaps. Extensive wind-tunnel studies, however, were made of more complex high-lift devices both before and during World...

graph of lift effect with various airfoil designs
Figure 5.3 - Effect of various types of high-lift devices on airfoil section maximum lift coefficient.


[108] War II. A summary of the state of the art of high-lift device design at the end of World War II is indicated in figure 5.3, in which the maximum lift capabilities of airfoils equipped with various types of leading and trailing-edge high-lift devices are shown. The maximum lift coefficient of an airfoil equipped with a plain flap, split flap, single-slotted flap, double-slotted flap, and double-slotted flap in combination with a leading-edge slat are shown in figure 5.3. The use of a double-slotted flap and leading-edge slat increases the maximum lift coefficient from about 1.4 for the plain airfoil to a value slightly over 3.2. The Douglas A-26 was the first aircraft to employ a double-slotted flap, and the combination of double-slotted flap and slat was not used to any great extent until well after World War II Many of today's jet transports employ double-slotted flaps or even triple-slotted flaps in combination with leading-edge slats and flaps. The leading-edge flap is not shown in figure 5.3 since it was a German development and was not known in this country until German data became available following the end of World War II. Many general aviation aircraft of today employ either plain flaps or single-slotted flaps. The airfoil with double-slotted flaps and slats shown at the top of figure 5.3 with a maximum lift coefficient of about 3.8 employed boundary-layer suction through a single mid-chord slot to delay separation of the boundary layer and thus increase the maximum lift coefficient. This concept was the subject of numerous experiments in wind tunnels but has never been utilized on a production aircraft. Various types of boundary-layer blowing have been employed for improving the maximum lift coefficient. This type of boundary-layer control became practical, however, only after the development of the turbine engine. The values of maximum lift coefficient given in figure 5.3 are for a two-dimensional airfoil section and are higher than would be obtained on a three-dimensional airplane wing equipped with partial span flaps.

Drag Cleanup
The internally braced monoplane with retractable landing gear, typified by the Douglas DC-3 shown in figure 4.12, would ideally be expected to have a zero-lift drag coefficient only slightly in excess of that which would be calculated with the use of the total wetted area of the airplane and a skin friction coefficient corresponding to a turbulent boundary layer. Such an ideal drag coefficient, however, is never achieved in actual service aircraft. The German Messerschmitt 109 fighter, for example, is shown in reference 72 to have a zero-lift drag [109] coefficient about twice the value corresponding to the ideal based on wetted area and a turbulent skin friction coefficient. The increases in drag above the ideal value result from one or more of the following:
(1) Projection of various items outside of the smooth basic contour of the aircraft
(2) Roughness or unevenness in the aircraft surface
(3) Unintentional leakage of air through the aircraft structure
(4) The use of large quantities of excess air for various cooling functions
(5) Areas of local flow separation
Experience gained during the 1930's from the investigation of full-scale aircraft in the Langley full-scale (30- by 60-foot) wind tunnel had given an indication of the importance of detailed design in the achievement of low drag coefficients on actual full-scale aircraft. Thus, during World War II, some 23 military aircraft were the subject of drag cleanup investigations in the Langley full-scale tunnel. Individual reports were issued following the investigation of each aircraft, and two separate summary reports covering the drag cleanup work were issued by the close of World War II. Recently, the data obtained during these various investigations have been summarized again and issued as a NASA publication, reference 42. The data obtained in the drag cleanup tests during World War II have been reissued in a modern report in order that they may be more available to the designers of modern general aviation aircraft.
A full-size aircraft installed in the Langley full-scale tunnel for a drag cleanup investigation is shown in figure 5.4. In this case, the aircraft is a Curtiss SB2C-4 Navy dive bomber popularly known as the Helldiver. The aircraft is mounted on three struts, two of which are located near the longitudinal center of gravity on either side of the aircraft center line and the third is located near the tail of the aircraft. These struts are attached to scales from which the lift, drag, and pitching moment can be measured. The two large four-bladed fans visible in the background of the photograph are connected to 4000-horsepower electric motors that provide the power necessary to drive the tunnel. The top speed of the tunnel is about 100 miles per hour. An indication of the size of the tunnel is shown by the man standing on the lip of the exit bell of the open throat test section of the tunnel. The Langley full-scale tunnel was first put into operation in 1931, has been continually used through the years since then, and is still in use at this time.

A Curtiss SB2C-4 inside of a wind tunnel
[110] Figure 5.4 - Curtiss SB2C-4 mounted in Langley full-scale tunnel for drag cleanup investigation. [NASA]

Drag cleanup investigations are still performed even today. A modern twin-engine general aviation aircraft is the most recent example of such an investigation. The procedure followed in a wartime drag cleanup study consisted of the following steps: First, the aircraft was examined in detail, those features suspected of causing unnecessary drag were identified, and necessary changes to eliminate the suspected unnecessary drag were planned. The airplane was then put in a faired and sealed condition in which all protrusions were either removed or carefully faired, all openings were closed, and all external leaks were sealed. The airplane was then returned to its service condition, item by item, and the drag was evaluated for each step. The procedure is illustrated by the results contained in figure 5.5 taken from reference 42, which shows the sources of drag for the Seversky XP-41 aircraft. The XP-41 airplane was very similar in appearance to the Seversky XP-35 shown in figure 4.14. Figure 5.5 shows that the aircraft drag was evaluated for 18 different conditions, which are indicated by sketches on the left-hand side of the figure and described on the right-hand side of the figure. The drag coefficient of the clean airplane was 0.0166, as....

Condition number


CD (CL = 0.15)

[Delta] CD

[Delta] CD, percent a


Completely faired condition, long nose fairing





Completely faired condition, bluntnose fairing





Original cowling added, no airflow through cowling





Landing-gear seals and fairing removed





Oil cooler installed





Canopy fairing removed





Carburetor air scoop added





Sanded walkway added





Ejector chute added





Exhaust stacks added





Intercooler added





Cowling exit opened





Accessory exit opened





Cowling fairing and seals removed





Cockpit ventilator opened





Cowling venturi installed





Blast tubes added





Antenna installed







a Percentages based on completely faired condition with long nose fairing.
digram of XP-41 drag sources
Figure 5.5 - Experimental study of drag sources on Seversky XP-41. [from ref. 42]


[112] ...compared with 0.0275 for the aircraft in the service condition. In order to convert the clean configuration into a useful practical aircraft, the drag was increased by about 65 percent of the value obtained for the clean aircraft. All the additional drag, however, was found to be unnecessary. Further tests and analyses showed that the additional drag could be reduced by more than one-half through careful tailoring of various aspects of the design. The drag coefficient of a practical service aircraft of the XP-41 type was accordingly reduced from 0.0275 to 0.0226. The data in figure 5.5 indicate that the increments in drag coefficient corresponding to the 18 steps of the cleanup process are generally rather small and, in many cases, only a few percent of the total drag coefficient. Yet, taken all together, these increments add up to an impressive total. Important performance improvements resulted from the drag cleanup of the 23 military aircraft in the Langley full-scale tunnel. In many cases, the gains associated with care and attention to detailed design were found to be greater than the differences in drag between airplanes of different configurations. The drag cleanup work made an important contribution to the refinement of high-performance propeller-driven aircraft during World War II, and the gains resulting from the program often spelled the difference in performance between victory and defeat in the air.

Compressibility Effects
Until the late 1930's, aircraft were designed on the assumption that the air flowing over the wings and other surfaces was essentially incompressible, like water. As speeds and altitudes increased, however, the effects of compressibility on the flow over the aircraft began to assume increased importance. The ratio of the aircraft speed to the speed of sound provides a useful index for gaging the speed at which significant compressibility effects begin to manifest themselves on a particular aircraft. This ratio is called the Mach number, in honor of the famous Austrian physicist Ernst Mach. The critical Mach number is defined as the aircraft flight Mach number at which the local Mach number over some portion of the aircraft, such as the upper surface of the wing for example, equals unity; that is, the flow at this point has reached sonic velocity.
Large changes in the pressures, forces, and moments acting on a wing or body occur at Mach numbers somewhat in excess of the critical value. These changes in aerodynamic characteristics result from the formation of shock waves and attendant flow separation behind the [113] shock wave. An example of the effects of compressibility on the lift and drag characteristics of a 15-percent- thick airfoil section is shown in figure 5.6 (from ref. 48). The section lift coefficient and the section drag coefficient are shown as a function of Mach number in figure 5.6(a) and (b), respectively. Precipitous reductions in lift coefficient occur with increases in Mach number beyond the critical value. The Mach number at which the lift begins to show a sharp decrease becomes smaller as the angle of attack is increased since the critical Mach number decreases with increasing angle of attack. Apparent also is the large reduction in lift-curve slope at the higher Mach numbers. For example, at a Mach number of 0.4, the lift coefficient increases from 0.2 to about 0.72 as the angle of attack varies from 0 to 5 whereas, at a Mach number of 0.8, increasing the angle of attack from 0 to 5 results in an increment in lift coefficient of only about 0.2. The drag coefficient shows a large increase with Mach number as the Mach number is increased beyond the critical value. For example, at an angle of attack of -1, the drag coefficient increases from about 0.0 15 at a Mach number of 0.65 to 0. 13 at a Mach number of 0.9.
Engine cowlings, canopies, propellers, fuselages, and other aircraft components were also found to be subject to large compressibility effects. Although not shown by the data in figure 5.6, large Mach number effects were found in the pitching-moment characteristic of the airfoil and in the effectiveness of various types of trailing-edge control surfaces. The effect on the airplane of these various changes in aerodynamic coefficients manifested itself in the form of a limiting speed, large changes in stability and trim characteristics of the aircraft, important reductions in the control power of the control surfaces, buffeting, loss in propulsive efficiency and various types of aircraft oscillation, and unintended maneuvers. In some cases, aircraft flown deep into the compressible regime became completely uncontrollable and could not be recovered. Loss of the aircraft and pilot frequently occurred under these circumstances. The state of understanding of compressibility effects in 1941 is outlined in reference 105, which was initially issued as a confidential report; a broader survey of knowledge in the field of compressibility aerodynamics is given in the Wright brothers lecture for 1944, which is cited as reference 106.
Extensive investigations were undertaken in the United States and Europe in an effort to better understand compressibility phenomena and, in particular, to devise design methods for increasing the value of the critical Mach number and reducing the adverse effects of compressibility that occur beyond this Mach number. These efforts were...

graph of lift coefficient and Mach
(a) Section lift coefficient.
graph of drag coefficient and Mach
(b) section drag coefficient.
[114] Figure 5.6 - Lift and drag characteristics of NACA 2315 airfoil section as function of Mach number for several angles of attack. [data from ref. 48]

[115] ...hampered by fundamental difficulties in both theoretical and experimental methods of investigation. The governing equations for flows near Mach number 1.0 proved intractable to closed-form solution. Adequate solutions to these nonlinear equations were not possible until the advent of the large-capacity, high-speed digital computer in the late 1960's and 1970's. Practical theoretical approaches to the compressibility problem during the war years usually involved the application of relatively simple correction factors to results obtained under the assumption of incompressible flow. These correction factors worked fairly well up to Mach numbers relatively close to the critical value but broke down completely at higher Mach numbers. The wind tunnel which had proved so useful in past aerodynamic investigations also became of questionable value at Mach numbers somewhat in excess of the critical value. At some Mach number, not too much higher than the critical value for the airfoil or body, the tunnel "choked," which meant that no higher free-stream Mach numbers could be obtained. A Mach number range between the subsonic choking value and some supersonic value, such as 1.2 or 1.3, was not available for wind-tunnel investigations. Supersonic tunnels operating beyond a Mach number of 1.2 or 1.3 were possible but were of little practical interest during the World War II time period. The solution to the problem of wind-tunnel choking was not found until the advent of the slotted and perforated- throat wind tunnel in the early 1950's.
In spite of these experimental and theoretical difficulties, a good deal of progress was made in devising improved configuration concepts for high-speed flight. The laminar-flow airfoil sections described previously did not achieve the desired objective of extensive laminar flow in flight; however, the pressure distributions of these airfoil shapes resulted in critical Mach numbers that were significantly higher for these sections than for other airfoil sections having the same thickness ratios. Most aircraft designed in the United States after 1940 employed the NACA laminar-flow airfoil sections or some modification of these sections, primarily because of the advantages they offered as a means for increasing the critical Mach number. The original NACA cowling, which was developed before aircraft speeds reached high enough values for compressibility effects to be important, had a critical speed of only about 300 miles per hour at 25 000 feet. New cowling shapes were developed that ultimately raised the critical speed to almost 600 miles per hour. Studies of various wing-body combinations led to configuration concepts that resulted in reduced interference effects and, hence, higher critical Mach numbers.
[116] Propellers usually encounter the adverse effects of compressibility at flight Mach numbers below that at which the aircraft configuration itself penetrates the critical region because portions of the blades of the propeller, particularly near the tip, are traveling at a higher speed relative to the air than the aircraft itself. Compressibility problems on aircraft propellers were first encountered during the 1930's, and research studies were made in those years in an effort to improve propeller design. This work continued on through World War II. One major investigation that gives an indication of the type of research undertaken in the development of improved propellers is described in reference 108. New planform shapes, new twist distributions, and new airfoil sections designed especially for propellers all combined to result in significant increases in the stream Mach number at which the propeller showed serious losses in efficiency. It seemed clear, however, that the propeller was likely to constitute the ultimate limitation on the speeds that could be reached with aircraft employing this means of propulsion.
The basic principles underlying the proper design of aircraft configurations intended for flight at high subsonic and transonic Mach numbers were fairly clear by the end of World War II The need for small thickness ratios on wings and tail surfaces and high fineness ratios on bodies became increasingly evident by 1945. The P-51D airplane, one of the best of the United States fighter aircraft of World War II, employed a wing of about 15-percent thickness ratio; by contrast, the wings of transonic and supersonic aircraft of today are more likely to be of the order of 4 to 5 percent in thickness ratio. The use of wing sweepback as a means for increasing the critical Mach number and reducing the adverse effects of compressibility beyond the critical Mach number was first proposed in the United States in 1945. (See ref. 77.) The advantages of sweepback had been recognized in Germany at an earlier date, and the Messerschmidt ME-163 tailless rocket fighter employed a sweptback wing. This aircraft saw limited operational use toward the end of World War II but was not particularly effective as a fighter because of the capricious nature of its rocket propulsion system. The use of wing sweepback, together with small thickness ratios and high fineness ratios, and later combined with the transonic area rule, provided the basic configuration elements needed for successful aircraft of high subsonic and transonic speed. The loss in propulsion efficiency at high subsonic Mach number remained the stumbling block to the development of successful aircraft for use at high subsonic and transonic speeds. The advent of jet propulsion solved this problem and, in addition, was capable of producing the large powers required [117] for flight at these high speeds with a simple and light type of propulsion system. The large power-producing characteristic of the turbine engine is related directly to the large air-handling capability of this engine as compared with the reciprocating engine. The jet engine then became the basis for all high- performance aircraft developed after about 1945. When used in combination with the configuration concepts just discussed, this propulsion system resulted in the high-performance subsonic and supersonic aircraft in operation today. Jet aircraft form the subject of part II of this book.
Flying and Handling Qualities
The flying and handling qualities of an aircraft have been of great interest since the earliest days of aviation. As pointed out in chapter 2, an aircraft with good handling characteristics must obey the pilot's inputs precisely, rapidly, and predictably without unwanted excursions or uncontrollable behavior and, finally, without excessive physical effort on the part of the pilot. Preferably, the aircraft should possess these desirable characteristics throughout its performance envelope. A well-known NACA test pilot of World War II and earlier years, Melvin N. Gough, put it in a slightly different form when he stated that "the flying qualities of an aircraft may be defined as the stability and control characteristics that have an important bearing on the safety of flight and on the pilot's impressions of the ease and precision with which the aircraft may be flown and maneuvered." For many years, there was considerable speculation as to what flying characteristics were desired in an airplane, and the entire subject was discussed in terms of the qualitative opinions of various pilots. Several years prior to World War II, a flight research program was undertaken in which the response characteristics of the aircraft following known control inputs were measured and correlated with pilots' opinions of the behavior of the aircraft, and, finally, related to the engineering parameters employed in the design of the aircraft. NACA continued the investigation of flying and handling qualities of various aircraft and, by the beginning of World War II, had assembled complete qualitative information on 12 different aircraft. From the fund of information accumulated in these tests, it was possible, in 1941, for NACA to prepare a set of requirements (ref. 53) for satisfactory flying qualities in terms of quantities that had been measured in flight and could be estimated by engineers during the design of a new aircraft.
[118] Flying qualities requirements may be listed under the broad headings of longitudinal stability and control characteristics, lateral stability and control characteristics, and stalling characteristics. The scope of flying qualities specification at that time is indicated in the following list of categories in which criteria were developed:
A. Requirements for longitudinal stability and control:
(1) Elevator control and takeoff
(2) Elevator control in steady flight
(3) Longitudinal trimming device
(4) Elevator control in accelerated flight
(5) Uncontrolled longitudinal motion
(6) Limits of trim due to power and flaps
(7) Elevator control and landing
B. Requirements for lateral stability and control:
(1) Aileron control characteristics
(2) Yaw due to ailerons
(3) Rudder and aileron trim devices
(4) Limits of rolling moment due to sideslip
(5) Rudder control characteristics
(6) Yawing moment due to sideslip
(7) Crosswind force characteristics
(8) Pitching moment due to sideslip
(9) Uncontrolled lateral and directional motion
C. Stalling characteristics:
(1) Pitching-moment characteristics
(2) Rolling- and yawing-moment characteristics
(3) Control forces
(4) Recovery
These various categories are not discussed in detail here and are only given to indicate the extent of design criteria available at that time. Most of the control criteria involved specification of the control power, that is, the ability of the control to cause the aircraft to respond in the desired manner, and control force and control-force gradients that relate to the physical effort the pilot must exert in order to actuate the controls by an amount needed to give the desired response. For example, the elevator control in accelerated flight is expressed in terms of the pounds of force that the pilot must exert on the control column in order to produce an acceleration of 1 g.
[119] The U.S. Army and Navy revised the general NACA flying qualities specifications to their immediate specific requirements and looked to NACA to continue its investigations and refinement of existing and new military aircraft. By the end of World War II, the total number of airplanes studied in flight by NACA increased from 12 to 60. A good discussion of the state of understanding of aircraft stability and flying qualities at the close of World War II is given in reference 102. The study and refinement of aircraft flying and handling qualities have continued through the years as aircraft speed, size, and configuration have changed and today form a highly sophisticated branch of aeronautical engineering.
Although not a specific part of the flying qualities requirements as defined in reference 53, aircraft spinning and spin recovery might be briefly mentioned under category C above, designated as stalling characteristics. In 1936, NACA put into operation at the Langley Memorial Aeronautical Laboratory the world's first vertical free-spinning wind tunnel. This tunnel was developed for the purpose of studying the control motions required to permit rapid and desirable recovery of an aircraft once it was in a spin, and for developing stability and control criteria for aircraft design so that the aircraft would have desirable spin recovery characteristics. During the war years, spin investigations were conducted in the free-spinning tunnel on approximately 150 different military airplane designs to determine recovery characteristics from developed spins. From the results of these investigations, criteria were developed for selection during the design process of proper design parameters so as to ensure good spin recovery.
Summary Comments
The preceding paragraphs describe four aspects of aerodynamic technology that were the subject of intensive research and refinement during World War II. These are intended to serve only as typical examples of the type of detailed research and refinement that took place in all technical areas involved in aeronautical engineering. Many other aspects of the science of aeronautics were under intensive investigation. Structures and materials technologies were advanced and methods of mass production were developed that resulted in the output of over 95 000 airplanes in the United States during one year of World War II. Propulsion technology, including engines, superchargers, fuels, and so forth, was the subject of intensive research and development. As an example, the magnificent Rolls-Royce Merlin engine developed about 970 [120] horsepower as installed in the original prototype of the Hurricane fighter; by the end of World War II, some versions of the Merlin engine developed 1600 to 1700 horsepower. The appearance of military aircraft changed very little in the time period between 1935 and 1945; however, the combat aircraft that existed in 1945 was far superior to its 1935 progenitor because of intensive work aimed at detailed refinement of all aspects of aeronautical design.