After World War II, the United States undertook a secret operation known as Project Paperclip, the aim of which was to exploit German scientists and engineers for American research, and to deny the technical expertise of these individuals to the Soviet Union. More than 1,600 scientists and their dependents were recruited and brought to the United States by Paperclip and its successor projects from the late 1940s through the early 1970s. Among these technical personnel were Drs. Wolfgang Noeggerath, Rudolf Edse and Gerhard Braun, who were employed in the Foreign Exploitation Section of the Air Documents Division, Intelligence, T-2, Wright Field beginning in 1945.
Dr. Braun was assigned to the Aircraft Laboratory, while Drs. Edse and Noeggerath were employed in the Rocket Laboratory. The author has uncovered little else about these engineers beyond their activities in Germany. Dr. Braun worked for the German Luftwaffe and was largely responsible for the creation of the “Hecht” and “Feuerlilie” anti-aircraft rocket projects at LFA Braunschweig. There, he directed activities concerning high-speed missiles with Drs. Edse and Noeggerath, who were responsible for propulsive details in addition to their other projects.
On October 15, 1945, the Army Air Force requested a design for an anti-aircraft rocket from Drs. Braun, Edse, and Noeggerath. Dr. Braun was primarily responsible for the organization and principle details of the project. The Army requested that the anti-aircraft missile incorporate the following performance characteristics:
|Velocity at 60,000 ft||Mach 2|
|Payload||100 – 200 lbs|
In response, the engineers issued their proposal report on December 27, 1945. The missile study, humorously labeled the CH-64 “Chow Hund” (Hund of course being the German word for “hound”), was quite advanced for its day and incorporated what had been learned in previous experiments with the Hecht, Feuerlilie and other German anti-aircraft missiles during World War II. The following is a summary of this rather technical but interesting report.
Dr. Edse prepared the analysis and comments concerning the booster rockets.
His personal experience and general theoretical investigations had shown that boosters were employed to advantage in anti-aircraft rockets. This was due to the fact that the missile received the kinetic energy of the boosters more or less without penalty, since the mass of the boosters did not require further lifting or accelerating after they were exhausted.
Inasmuch as the rocket was to fly at supersonic speeds, it appeared practical to install boosters sufficiently powerful that sonic speed was attained by the time the boosters dropped. In this manner, the difficulties of achieving an aerodynamic form suitable for both supersonic and subsonic speeds were avoided; other known solutions of this problem involved a compromise of supersonic aerodynamic characteristics. It was therefore possible to choose dimensions especially for the supersonic speed range since negligible stability difficulties were encountered at subsonic speeds due to the high accelerations during takeoff.
A high initial acceleration shortened the time between takeoff and the strike at the target, providing a tactical advantage. Acceleration was not of such magnitude, however, that the stress limitations of the power plant and steering apparatus was exceeded, or that a weight penalty was introduced by very high combustion pressures in the boosters, which would have necessitated greater wall strength.
For these reasons, four powder boosters with identical combustion chambers developing a total thrust of 48,000 kg (105,600 lbs) for a combustion period of three seconds were decided upon. Thus, the total impulse of the booster rockets was equal to 3 x 4 x 12,000 = 144,000 kg/sec (317,000 lb/sec). The booster rockets would have burned 720 kg (1,587 lbs) of powder. The burning time of the boosters were not of long duration, since steering was practically impossible until the boosters were jettisoned and since the boosters were to have been dropped from as low an altitude as possible. The altitude chosen for the booster release was approximately 500 m (1,640 ft). The boosters were attached to the rocket fuselage in a similar arrangement to the ones used on the Feuerlilie and “Enzian” rocket projects in Germany. If attached behind the rocket, the boosters almost certainly wouldn’t have damaged the main rocket when released, but tandem mounting had the great disadvantage that the center of gravity was shifted towards the rear, necessitating exceptionally large control surfaces.
In order to achieve light weight, the walls of the combustion chamber had to be relatively thin, and combustion had to take place at low pressure. The full benefits from this weight saving could not be completely realized, however, since the thrust necessarily decreased with a reduction in the rate of expansion.
Experiments were to have determined the type of powder used. Double Base Powder (Nitroglycerine – Nitrocellulose), Di-nitro Di-glycol Powder, and Penta-Erithrite-Tetra-Nitrate were to have all been investigated; such powders were considered of great importance for the future development of rockets. It was also intended to further develop powders that possessed extremely high combustion velocities in a low pressure range, which would permit the use of plastic materials for the construction of rocket combustion chambers, eliminating the need for a nozzle.
Certain difficulties could have also resulted from uneven combustion or stalling of one or more of the boosters. This was avoided in the Feuerlilie by connecting the four boosters with equalizing pipes. Since such an arrangement was impractical for the CH-64, the booster nozzles were inclined towards the axis of the fuselage, so that the thrust of each individual booster was directed though the center of gravity of the missile. Thus, the failure of a booster did not cause a moment around the center of gravity but only a component force resulting small lateral movement only.
In order to minimize damage, the jettisoned boosters were to have descended by means of parachutes. Another way of reducing the damage caused by falling boosters was to break up the powder chamber in the air by blasting it into small pieces. This was difficult to accomplish, however, since the booster material was too tough to be readily broken up, and since any blasting charge placed around the outside of the booster exploded before the end of combustion due to the high temperature developed in the combustion chamber (800 degrees Centigrade on the Feuerlilie).
|Overall Length with Nozzle||4,500 mm =||….14.8 ft|
|Outside Diameter||250 mm =||….9.83 in|
|Length of Nozzle||200 mm =||….7.87 in|
|Weight of Chambers||250 kg =||….551 lbs|
|Weight of Charge||200 kg =||….442 lbs|
|Casing and Parachute||50 kg =||….110 lbs|
|Thrust||12,000 kg =||….26,455 lbs|
|Combustion Period||3 seconds =|
The weight of the experimental boosters to be used repeatedly in experiments was to have been increased to 350 kg (771 lbs) to insure sufficient structural strength to withstand experimental abuse. Each booster was fastened to the main rocket at two points. The front fastening was located immediately aft of the helium bottles and served to transmit longitudinal thrust to the fuselage. The rear fastening carried only lateral forces. This method of mounting had the advantage that the nitric acid tank located to the rear of the front support was placed in tension by the longitudinal thrust of the boosters, which eliminated buckling tendencies and made lighter construction possible. The structural components forward of the front fastening, on the other hand, had to demonstrate great stiffness for aerodynamic and other reasons.
Jettisoning of the boosters was accomplished by the explosion of a powder charge installed in the front and rear booster mounting assemblies. A negligible lateral impulse was caused by the explosion since opposite pairs of boosters were thrown off simultaneously. The powder charges in both cylinders were of such strength and proportions that the boosters were thrown free of the main rocket without damaging fuselage or fins. Tests on the “Enzian” had shown that such a method was feasible. Size of the charges was to have been determined with mock-ups preliminary to ground tests.
Launching was to have been accomplished from a device that, for practical purposes, could have been a mass-produced gun carriage, using launching rails instead of a gun barrel. This would have provided a cheap, sturdy, mobile undercarriage capable of adjustment in the vertical and horizontal planes. A gun carriage remotely controlled by a radar device would have been best. For high-speed targets, mechanical computing devices appeared to be practical and even necessary. For high initial acceleration, the length of the rails could have been comparatively short (12-15 m or 40-49 ft)
Power Plant and Fuel
Dr. Noeggerath worked out the proposal for the rocket motor, work program, and essential requirements for development of the “Chow Hund” propulsion unit. The power plant of the main rocket was designed to deliver a thrust of 6,000 kg (13,230 lbs.) at 5,000 m (16,400 ft). This value was to have been smaller at lower altitudes and greater at higher altitudes. The rocket power plant had the following characteristics:
Oxidant: 98% HNO3, specific gravity 1.51
Fuel: Ergol (self-igniting with nitric acid), specific gravity approximately .97
Fuel Pressurization: Pressurized with helium
Ignition: Not required – use of self-reactant fuels was contemplated
Pressure in combustion chamber: 25 atmospheres
Thrust: 6,000 kg (13,230 lbs) at 5,000 meters (16,400 ft)
Mixing ration of oxidant to fuel: 3.8 : 1 by weight (assumed value)
Total weight dry: 813 kg = 1,792 lbs
Fuel and pressurization weight: 1,107 kg = 2,440 lbs
Total weight wet: 1,920 kg = 4,233 lbs
Length: 5,841 mm = 19.3 ft
Diameter: 640 mm = 2.1 ft
Volume: 1.88 cu m = 66,400 cu ft
Specific weight of power plant (wet): 1.02 kg/ltr = 63.5 lbs/cu ft
Specific weight/unit momentum: .00915 lbs/lb sec
Specific volume/unit momentum: 8.98 ltr/ton sec = .288 cu ft/ton sec
Delivery of fuel to the combustion chamber was accomplished by pressurized helium, which was contained in two spherical tanks at a pressure of 220 atmospheres. A weight saving of 42 kg (92.5 lbs) was realized by using helium instead of nitrogen. In order to insure an uninterrupted supply of propellant fluids during transverse accelerations, the tank risers, or standpipes, were equipped with a flexible “trunk.” The free end of this trunk, being subject to the same accelerations as the fluid, would have been always immersed. The transfer pipes that conducted the pressurized helium to the fuel tanks were of the same cross-sectional area as those carrying fuel to the combustion chamber. They were arranged in two symmetrical groups outside the skin, and were covered by a sheet metal hood.
Operation was started by electric detonation of an explosive protection valve located between the helium tank and the pressure reduction valve, which admitted helium to the fuel tanks at a pressure of 35 atmospheres. The liquids were forced through the standpipes and transfer pipes in the mixing nozzle, where spontaneous ignition occurred. The HNO3 was routed through the cooling jacket where it served to hold the temperature of the discharge nozzle and combustion chamber within proper limits before it was admitted to the mixing nozzle. At the points where the helium entered the fuel tanks, cheek valves were installed to prevent any transfer of fluids between tanks after combustion had ceased. Adequate protection against the mixing of propellant liquids during storage was provided by installing blow valves in all fuel outlet lines, and between the helium pressure reduction valve and the tanks. These valves completely isolated each tank and were designed to rupture at a pressure difference of approximately 5 atmospheres. For test purposes other necessary valves could have been installed so that the tests could have been stopped at any desire point.
Axial thrust of the boosters was applied to the rocket forward of the nitric acid tank, which stressed the entire power plant in tension during the initial acceleration period. This arrangement was necessary to avoid undue structural weight, since stress values were approximately ten times as large at that time as during subsequent flight. Lateral forces from the booster rockets were equalized by a transverse tubular spar which passed through the fuselage between the mixing nozzle and Ergol tank. To eliminate the possibility of failure by buckling during the second acceleration period, when the power plant was subject to compression, the Ergol tank was reinforced by wing structure projections and the HNO3 was strengthened by a layer of corrugated sheet metal.
Tank construction was of high grade, weldable steel. The tanks were to have been hydrostatically tested to pressures of 320 atmospheres for the helium tanks and 50 atmospheres for the propellant tanks, which induced a tensile stress of 3,800 kg/cm3 (54,000 psi) in the material. The HNO3 tank was protected against acid corrosion by an internal lining of pure aluminum. The combustion chamber and discharge nozzle were made of steel, the mixing nozzle of light alloys and steel, the piping of light metal, and accessories of light alloys and stainless steel.
A decision had to be made as to whether the apparatus was intended to be expendable, since this consideration greatly influenced construction design. Full operating requirements were to have been made available, such as the temperature in which launching was to occur, length of time the apparatus was to be stored ready for firing, etc.
Design was also greatly influenced by fuel consumption and the weight ratio of oxidant and fuel, but these values could only have been determined by tests of full scale motors. Items such as the mixing nozzle, combustion chamber, exhaust nozzle, and individual accessories were to have been developed and ultimately perfected by basic preliminary tests on a small scale with later supplementary full-scale tests. To expedite this work, Dr. Noeggerath proposed making these tests in three steps:
- Preliminary tests with small combustion chambers of 200 kg (441 lbs) thrust, to determine the optimum mixing nozzle system and to obtain construction data for a combustion chamber and exhaust nozzle where adapted to the proposed fuel system.
- Tests with a medium size motor of approximately 2,000 kg (4,140 lbs) thrust.
- Tests with the full-scale motor of 6,000 kg (13,230 lbs) thrust.
The reliable operation of the power plant and the rate of consumption depended essentially on the design of a mixing nozzle for the particular fuel to be employed. On the basis of existing knowledge, and especially in view of the possibilities existing in the US to improve fuels, it was assumed that such development would proceed rapidly and yield valuable results.
The exhaust nozzle, the shape of which influence fuel consumption considerably, could be correctly dimensioned only after the combustion characteristics of the particular fuel decided upon had been determined by tests.
In general, accessories such as expansion valves, safety valves, fuel distributor parts, inlets with vents, flexible standpipes, and pressure reduction valves had to be specially developed, since the design requirements for rockets were difficult to achieve and often contradictory. In 1945, relatively few rocket engines had been completely developed, and the available data merely served to establish general principles of construction and to indicate that accessories had to be designed and tested for each individual case. The construction of the pressure reduction valve had to be best assigned to the appropriate firm.
The tanks could have been constructed before the exact mixing ratio had been established since subsequent changes in volume had to be made by altering tank length, which would not have affected construction. The details of construction and the manufacture were to have been assigned to a firm specializing in this type of work.
An ergol, such as had been developed by the LFM in Germany, was proposed for a fuel. Due to the fact that the reactions of the ergols greatly depended upon their chemical composition, on the amount and type of chemical impurities present in the primary substances used to produce them, and since the primary materials available in the US were probably not chemically identical to those used in Germany, it would have been necessary to subject fuels developed in the US to further investigation and experimentation. Since conditions were very favorable in the US for obtaining fuels that were not available in Germany, and since it could be expected that further improvements in the ergols could be achieved by minor changes in their composition, it seemed justified to conduct further laboratory tests before final selection of the fuel was made.
To obtain data useful to future projects, a thorough evaluation of tests conducted in connection with the development of this missile was to have been made. In order to establish basic principles for this purpose, however, extensive thermodynamic calculations had to be first performed. On the basis of these considerations, a work program for the development of the CH-64 rocket was proposed.
The size of the rocket (weight and volume of the fuselage) was a function of the required altitude, speed, and specific consumption of the propellants of the power plant.
The overall aerodynamic and structural data for the proposed Chow Hund CH-64 anti-aircraft rocket were as follows:
|Take-off weight with boosters||4,500 kg||(9,920 lbs)|
|Take-off weight, boosters jettisoned||2,400 kg||(5,290 lbs)|
|Powder weight, boosters||800 kg||(1,764 lbs)|
|Fuel||1,100 kg||(2,425 lbs)|
|Warhead||100 kg||(220 lbs)|
|Length||7.5 m||(24.6 ft)|
|Diameter||.64 m||(2.1 ft)|
|Volume||2.0 cu m||(70.6 cu ft)|
|Maximum thrust of boosters||48,000 kg||(105,820 lbs)|
|Booster combusion period||3 sec|
|Maximum thrust of rocket||6,000||(13,230 lbs)|
|Combustion Period||35 sec|
|Velocity at 60,000 ft||800 m/sec||(2,630 ft/sec or 2.7 Mach)|
|Approximate ceiling||140,000 ft|
|Range, in excess of||45 miles|
Anticipated methods for direction control largely determined the choice of number of fins or wings. A two wing design (monoplane) possessed the advantage of smaller drag and wing weight, and, for a specified total wing weight; exhibited a smaller turning radius than the four wing counterpart since the wing area of the former, in the plane of the wings, was greater. Since the type of control was not specified, and since two-winged missiles were somewhat more difficult to steer, a cross-winged (four wing) construction was chosen, despite the aerodynamic disadvantages.
For cross-winged construction, control was accomplished by two methods. The first required that the lateral axis of the rocket remained horizontal during all flight attitudes, which necessitated the incorporation of a stabilizing device into the missile to eliminate any rolling motion about the longitudinal axis. With this arrangement one set of elevator surfaces at all times controlled pitch and the other lateral motion. Thus, if the ground control device were so designed that its manipulation and effect on the missile were similar to a conventional airplane control stick, fore and aft motion of the “stick” actuated only the set of elevators on the missile controlling pitch, and sideways motion actuated the set of elevators controlling lateral motion, or turning. If, on the other hand, the missile were permitted to roll during flight (roll being induced by any slight unsymmetrical irregularities of the aerodynamic shape), a device was required to coordinate the movements of the ground control “stick” with the movements of either or both of the sets of elevators for any instantaneous angular position of the missile. In this case a fore and aft motion of the ground control “stick” caused a motion of either elevator separately, or both together in the proper proportion. Control surfaces were actuated by a gyroscope in the rocket, or by polarized radio wave transmission with high frequency instruments either built into the rocket or on the ground.
The rocket was symmetrical in both the horizontal and vertical planes (planes intersecting on the longitudinal axis of the rocket and at right angles to each other) since the “upper” and “under” side of the rocket had identical aerodynamic functions and since lateral motion in a given direction was equivalent to lateral motion in any other direction during vertical flight. The requirement for high speed necessitated a slender fuselage (fineness ration 1 : 11.75) and small wings since the drag had to be held to a minimum to keep the power plant small. Small wing surfaces resulted in large radii of turn in the trajectory, but this did not appear to be a disadvantage since at such high velocities it was not possible to change the trajectory sharply enough to accomplish pronounced maneuvers in pursuit of a target. Furthermore, since this rocket was designed for essentially vertical flight, it was felt that controllable horizontal or descending flight was impossible, for various technical reasons. A miss would have always resulted, therefore, even with larger wings, if the rocket reached the altitude of its objective without demolishing it. It appeared most practical to aim the missile by means of an automatic computing device, and then to correct the flight path by remote control till collision with the target occurred.
Small wings were also desirable for strength considerations. If larger wings were used to achieve smaller radii of turn, air loading on the wings became so great that the whole rocket fuselage was heavily stressed, which required heavier construction with greater weight. A bi-convex airfoil section of 8% thickness was selected for the wing section, but a double-wedge section of the same thickness was also considered. In case the drag of the latter was shown to be only slightly larger, it was to have been given preference because of simplicity of construction.
The fuselage was a body of revolution with cylindrical midsection of 640 mm (25.2 in) diameter. The tip of the fuselage had an ogival shape, the radius of the curvature being equal to ten times the fuselage diameter. At a distance of approximately 650 mm (25.6 in) from the rear end of the rocket the fuselage tapered sufficiently to enclose the nozzle.
Since this missile was designed to operate through a wide range of velocities, considerable movement of the center of pressure was taken into consideration, which necessitated a very effective empennage. At the time, very little information was available about downwash in the supersonic range. It was safely assumed, however, that the empennage would have created less disturbance than by the wings since the wings constituted the larger assembly. For this reason, a canard type arrangement was selected (i.e., empennage located forward of the wing).
It was anticipated that considerable difficulty would have been encountered in the design of control surface actuating mechanisms since it was probable that very large hinge moments would have had to of been overcome. Wind tunnel and flight tests would have been required to develop an empennage shape with a relatively stable center of pressure. If such a shape was found, then the whole empennage was to have been installed as a rudder, exerting force around axis of rotation that passed through the center of pressure. If such shape was not found, controls of the type designed by Professor D.H. Wagner for subsonic speeds were to have been investigated for possible application to the missile. Dr. Wagner’s control had the advantage of small hinge moments and could have proven useful at supersonic speeds.
Because of the limited knowledge regarding supersonic speeds, it was possible to calculate only a rough approximation of static stability in that region. The calculated value indicated great static stability. Control surface deflections of approximately 1 ¾ degrees were shown to produce a change in angle of attack of one degree, which was believed adequate for satisfactory control.
For a flight performance estimate, drag surface for incompressible flow was assumed to be .082 square meters. Since very little data regarding skin friction at supersonic speeds were available, the calculations were only approximate. Fuselage drag was estimated from projectile data, the wind drag was obtained from airfoil measurements, and a correction was added for interference effects. At 60,000 ft, climbing speed was estimated to be 798 m/sec (2,625 ft/sec). Since sonic speed at this altitude was 293 m/sec (965 ft/sec), this velocity corresponded to a Mach number of 2.7, amply fulfilling performance requirements. Maximum altitude was roughly estimated to be 140,000 ft, or 24.5 miles. If it was assumed that the rocket descended with a gliding angle equivalent to 1 : 1.7 after attaining this altitude, it would have achieved a range of 45 miles. This range could have been great exceeded, however, since the best gliding angle was flatter than 1 : 1.7. Range could have been further increased by launching the rocket in other than a vertical direction in order to utilize the lift of the wings.
Wind Tunnel Tests
Aerodynamics as a science of continuities deals with partial derivatives that determine flow; these are of the fourth order and are not linear. A prediction of the aerodynamic loads and their distribution was only possible, therefore, when a new design was similar to the one whose characteristics were already known. The CH-64 was so dissimilar from designs of other anti-aircraft rockets, due to performance requirements, that it was impossible to calculate an exact performance mathematically. Extensive wind tunnel tests were considered necessary. These tests were to have not only furnished information as to the resultant aerodynamic loads and moments about the complete model, but would have indicated the distribution of the load between the individual components of the rocket, and the possibilities for improving the design. Data for solving the above problems were to have been made available by a comprehensive wind tunnel program.
Division of Labor
Although Dr. Braun desired to work on the development of an anti-aircraft rocket as an aerodynamicist and mechanical engineer, he stated that he would have only been able to work on a small part of such a complicated project. His proposed contribution to this work was as follows:
- Performance and supervision of wind tunnel tests.
- Evaluation of combustion chamber and wind tunnel tests for trajectory computations.
- Trajectory computations to determine the most suitable stabilizing apparata and controls, expected dispersion due to wind, variation in the combustion time and take-off weight, effective area of fire of the missile, and to establish a basis for an automatic lead computer.
- Collaboration with the group responsible for construction, in order to assure incorporation of desired aerodynamic characteristics and to supply the aerodynamic basis for stress analysis.
- Accomplishment of tests to achieve satisfactory jettisoning of boosters.
- Performance of flight tests and launching experiments to determine the aerodynamic characteristics of the rocket at high velocity.
To accomplish this work within a reasonable time, the personnel requirements would have included eight computers, two laboratory technicians, two construction engineers, and one aeronautical engineer.
Dr. Braun was not a specialist on the following unsolved problems:
- Development of boosters and the rocket motor.
- Construction of rocket structure and launching devices.
- Investigation of explosion effects in the air to determine effective blast area and blast pressure distribution.
- Development of mechanical and electronic controls.
- Stress analysis.
- Investigations to determine most satisfactory control methods, using models.
These problems were left to Drs. Edse, Noeggerath, and other specialists to solve.
Despite the advanced features of the CH-64, the Army Air Force elected not to undertake development of it. The author’s research has thus far not uncovered the exact reasons for its rejection, though one can make several reasonable speculations. It may have been a case of “too much, too soon,” in which the proposal was deemed too ambitious or risky. The existence of other AAM programs, such as the Nike missile, could have also been a deciding factor. The fact that the proposal originated from a group of unknown foreign scientists in a relatively obscure part of the AAF, and not a major aircraft manufacturer, also probably didn’t help the CH-64’s chances. Or it could have been the most common reason – lack of funding.
The CH-64 Chow Hund proposal remains an interesting example of German know-how being applied to an American security problem, i.e. the future vulnerability of its territory and forces to high speed strategic bombers. The proposal is also compelling because of its detailed description of contemporary German missile design and construction practices, which were then at the leading edge of the field. Though the CH-64 was turned down for development, it seems likely that the considerable expertise of Drs. Wolfgang Noeggerath, Rudolf Edse and Gerhard Braun was subsequently put to use on other programs. If any readers have more information on the careers of these German engineers, please contact the author.
All images from NARA Archives II, College Park, MD, RG 18
Dr. Wolfgang Noeggerath, et al, “Preliminary Proposal for a Guided Missile,” Intelligence T-2, Air Documents Division, Foreign Exploitation Division, Air Technical Service Command, Wright Field, Dayton, Ohio, December 27, 1945, in the files of the National Archives II at College Park, MD, RG 18.