The Hermes Project
The scale of the HERMES program was such that it studied every phase of guided missile work except large-scale production and warhead and fuse development. Research and development were conducted in missile-borne guidance and control ground launching and handling equipment flight test instrumentation, subsonic and supersonic ramjet engines, fuels and propellant combinations, and rocket engines.
Army Ordnance established the HERMES Project in 1945 after having recommended in 1944 a development project for a long-range rocket and the necessary handling equipment. General Electric Co. was selected to carry out the diversified program because its personnel had significant technical experience and facilities. At its peak of operation in 1952, the HERMES Project directly employed some 1,250 people.
At the time of the HERMES project’s termination on 31 December 1954, the principal accomplishments of the effort were:
(1) Development and successful flight testing of the highest performance large-size liquid fuel rocket in the United States to that date
(2) Development and successful testing of the largest solid propellant rocket motor flown to that date
(3) Development and successful testing in flight of the first stabilized platform, inertial guidance equipment in a ballistic missile
(4) Development and successful demonstration during flight of a portable, highly accurate, highly complex phase comparison radar, including command guidance functions
(5) Demonstration of optimum air-frame structural design
(6) Establishment of altitude and velocity records in the Bumper program (q.v.)
(7) Creation of a coded command guidance radar link
(8) Development and testing of advanced fuels and motors
(9) Development through static testing phase of a ramjet engine to operate at a speed of Mach 4
Utilization of HERMES development work was purely speculative in 1955; one possibility called for use of the HERMES inertial guidance system, after adaptation, in the Sergeant missile, but such a decision rested with the Sergeant contractor.
“Ordnance Guided Missile and Rocket Programs-HERMES” suggests the division of the many areas covered by the HERMES Project into three general categories: the A3 series of missiles; the A1 and A2 series of missiles; other missiles and supporting research. Included in the latter category are the V-2 and BUMPER series, covered in the preceding section. This chapter on HERMES Project will utilize the general breakdown suggested above. Discussion of the V-2 and BUMPER missiles will be incidental.
Other Missiles and Supporting Research
Early research and development work under the HERMES Project consisted of general study, component development and experimental work at White Sands Proving Ground with modified German V-2 material. HERMES V-2 activity continued until mid-1951. The BUMPER Program gave convincing demonstration of the feasibility of liquid-fuel propelled two-stage rockets. This series also yielded aerodynamic heating data. HERMES B began in 1946 as an investigation of ramjet propulsion with free-flight tests in 1947 utilizing the V-2 to carry an experimental ramjet diffuser to supersonic speeds. HERMES B1 was designed as an interim test missile to meet the May 1947 requirements from the Office, Chief of Ordnance, for design, development, and testing of a supersonic ramjet missile to carry a 5,000-pound warhead a minimum distance of 1,500 nautical miles at speeds of about Mach 4.
HERMES C formed the basis for the present Redstone Missile System although this relationship since has become dim.
HERMES II work covered development of a two-stage missile, the second stage to be a ramjet vehicle launched after boosts by the first stage, a modified German V-2 rocket. The second stage ram was 214 inches long, a little more than 50 inches in diameter; the wingspan was 183.86 inches, the rudder spanned 58.87 inches, and the elevator spanned 59.35 inches. Power came from two ramjet motors with rectangular cells with the ram openings placed at the lead edge of the wings. Maximum altitude was envisioned as 66,000 feet and speed as 3,180 feet per second. “Missile O,” first test vehicle with its V- 2 booster, was fired 29 May 1947 – this was the ill-starred missile that impacted in Juarez, Mexico. Since the modified V-2 booster served as the power source for this test firing, a description of the difficulties encountered and its impact is contained in the chapter concerning the V-2.
Three more HERMES II missiles, essentially the modified V-2 with a dummy ram wing mounted on the nose, were fired at White Sands Proving Ground 1949 and 1950. The objectives of these tests were to study the booster guidance and control system and to determine techniques for expulsion of the second stage vehicle. No information is available on the first firings of 3 January 1949 (also listed as 13 January and 14 January 1949 in various firing charts), 6 October 1949 and 9 November 1950 except that the latter proved successful in every aspect. Three more HERMES II missiles, essentially the modified V-2 with a dummy ram wing mounted on the nose, were fired at White Sands Proving Ground 1949 and 1950. The objectives of these tests were to study the booster guidance and control system and to determine techniques for expulsion of the second stage vehicle.
No information is available on the first firings of 3 January 1949 (also listed as 13 January and 14 January 1949 in various firing charts), 6 October 1949 and 9 November 1950 except that the latter proved successful in every aspect.
Prior to this last firing, HERMES II was discontinued as a tactical missile but continued as a ramjet test development, HERMES B was transferred from HERMES I to HERMES II and, on 31 October 1951, responsibility for HERMES II activity was transferred to the Ordnance Guided Missile Center at Redstone Arsenal with HERMES II re-designated the RVA-A-3 test vehicle. With the November 1950 test firing, all emphasis on the ramjet development program was to be continued in ground-based facilities with simulation of actual flight conditions. Ramjet development was continued at Malta test Station, N.Y., under the supervision of Redstone Arsenal until all Army ramjet work was terminated at the end of 1953.
Hermes A1 and A2 Series
Work on HERMES A1 as an anti-aircraft missile began in 1946, but the scope of the entire HERMES Project changed in 1947, limiting it to surface-to-surface missiles. Six A1 test vehicles were approved for guidance and control tests against surface targets; one of these was damaged beyond repair in static testing and the other five were flight tested between 19 May 1950 and 25 April 1951. Since guidance problems were of major interest in the A1 development program, project managers adopted the German “Wasserfall” missile configuration. This step also allowed HERMES Project to utilize the experience of the German guided missile personnel who had been assigned to HERMES Project at the inception of the program. HERMES motor experiments premised better performance with safer fuels, so the “Wasserfall” propulsion system was not incorporated into the A1 rocket.
The missile was 25.5 feet in length and 34.5 inches in diameter. Its control and stabilizing surfaces were four fixed wings and four controlled fins; its new liquid propellant power plant generated a thrust of 16,000 pounds to give the vehicle a maximum velocity of 1,800 miles per hour. Guidance controls were roll-and-pitch stabilization by gyro plus ground-originated radar signals operating jet and air control vanes. The rocket was highly maneuverable and an ideal test vehicle for this control experimentation.
Ground guidance was essentially a single tracking radar and a computer with the missile carrying a combination beacon-transponder. An SCR-584 radar was modified for tracking and command transmission, but an MPQ-12 radar, which was an improved version of the former model, was tested at White Sands Proving Ground after the A1 flight tests were completed.
A comprehensive set of electrical test equipment for HERMES A1 was designed by General Electric. This equipment resulted from experience gained in V-2 and BUMPER operations, which indicated that electrical failure of one type or another produced the most lost time in pre-firing checkout of the missile. A servo amplifier static test unit was designed for field use when control system simulators were not available. A pre-launch test unit performed propulsion system checks to indicate the presence of squib operating voltages and to simulate vehicle responses such as pressure switch operation.
Discontinuation of HERMES A1 as a tactical missile was ordered on 18 May 1950, the day before the first firing of the A1 at White Sands Proving Ground. The tactical design at that time was represented by HERMES A1E1, a 25-foot-high surface-to-surface missile with a 1,450-pound payload. HERMES A1E2, a modification of A1E1, was also designed as a tactical weapon; it stood four feet higher than the earlier missile, but its payload was identical. Fabrication of 12 A1E1 missiles was authorized on 10 November 1950. The Joint Chiefs of Staff then approved HERMES A1 type missiles for development as an interim tactical missile and fabrication of 12 A1E2 missiles was ordered on 4 January 1951. The A1E1 was canceled by the Office, Chief of Ordnance, on 23 October 1952; A1E2 work had been terminated in April 1952 due to the absence of a practical requirement.
HERMES A1 was first fired at WSPG on 19 May 1950 at 1100 hours. There was a delay of two hours because due to a sticking check valve in the oxygen fill line detected prior to launching. Thrust was lost within 10 to 20 seconds after launch; impact occurred at 39 seconds with a maximum altitude of 2,980 feet and a maximum velocity measured at approximately 300 feet per second. The missile’s point of impact was 6,390 feet north and 540 feet east of the target. Data records indicated three separate failures occurring during flight: the oxygen emergency dump valve opened about one second before launch causing the helium used to pressurize propellant tanks to be expended at four times the normal rate; the main alcohol start valve opened prematurely and in turn actuated the main oxygen start valve so that the short preliminary stage intended to facilitate firing did not occur; yellow flame in the jet about two seconds after take-off indicated burning iron and a small hole was found to have been burned through the inner wall of the rocket motor, although no variation in combustion pressure or acceleration was notes as a result of this failure. Command guidance systems seemed to work normally until impact; the vehicle was stable over the entire flight with only small deviation recorded in roll and yaw.
Round #2 was flight tested on 14 September 1950. The missile began to roll six or seven seconds after takeoff, and it rolled at a constant rate of 3.6 degrees per second. Despite the roll, the missile operated stably for approximately 41 seconds; complete loss of control occurred when the jet flame burned through the hydraulic servo covers. An explosion occurred 81 seconds in. The cause was believed to have been the accidental detonation of the shaped charge. Total elapsed time from launch to impact, maximum altitude and velocity, and range at impact are not available.
Round #3 did not get off the launch pad. The main stage burning did not take place as scheduled; “Emergency” and “Helium-Dump” switches were thrown and water was called for to extinguish the after burning. The vehicle was disarmed and liquid oxygen was allowed to bail off. Upon examination of the missile at the General Electric plant at Malta Test Station, the cause of the misfire was attributed to metallic fouling of the alcohol feed line with the original cause a failure of the cadmium plating in the alcohol jacket of the motor. Repairs were made and this missile was launched at WSPG in 1951.
Firing dates in 1951 were 2 February, 15 March, and 26 April, one of these missiles being the repaired #3 which did not function in its scheduled 28 September 1950 test. No information on the results of these specific firings is available. However, “Ordnance Guided Missile and Rocket Programs-HERMES,” page 23 states: “Although none of the A1 missile flights were completely successful, the functional operability of the system was demonstrated.”
The evolution of HERMES A1 into A1E1 and A1E2 ended with one A1E1 fabricated but not flown and no A1E2 configurations were even fabricated.
A version of the A1 missile-borne command guidance receiver decoder was adopted by the CORPORAL Project; pulse coding was employed in the command link to cut down on countermeasures or accidental interference.
HERMES A2 originated in 1946 as a wingless version of the A1, but the plan was not pursued. Studies in 1949, directed toward development of low-cost missile system, were labeled A2. A solid propellant motor system was selected and applicable portions of the A1 guidance system were to be employed. A 1,500-pound payload was to be carried to a 75-mile range. Development on HERMES A2. was confined to work on the propulsion system and the A2. motor was test-fired at the Air Force Missile Test Center in Florida using the RV-A-10 test vehicle; four firings were conducted. A hybrid propulsion system also was developed with polyethylene plastic as the fuel and concentrated hydrogen peroxide as the oxidizer. Hybrid motors of 1,050 and 20,000-pounds of thrust were repeatedly static fired, but the hybrid work was terminated in 1953 because due to a lack of a practical requirement.
HERMES A3 Series
Thinking in 1947 indicated the A3A would be slightly smaller than the V2; its propulsion system would provide 30,000-pounds of thrust. Twelve A3AE1 missiles, the original A3A design, were ordered on 10 November 1950 to be tested; subsequent A3A missiles, or A3AE2, were to carry the Type B 1,500-pound atomic warhead.
Requirements in September 1953 for a missile capable of carrying a 3,000-pound atomic warhead brought about the A3B, essentially a hurried redesign of the A3A. The A3A motor was adapted to the new missile with thrust specifications changed from 18,000 pounds to 22,000 pounds. A3B height and diameter measurements were 33 feet and 47 inches compared to 29 feet and 40 inches for the A3A. The guidance system was identical to that employed in A3A. In silhouette, the missile configurations differed thusly; A3A carried its bulk-maximum diameter-in the lower half of the fuselage while the apparent bulk in A3B was located in the forward half of the fuselage; tail fin surfaces on both missiles were triangular.
Trajectory and scheme of operations were the same for both missiles. Launch was vertical; the missile made a 1G turn four seconds after launch and then was guided by radio command along a straight line until radio cut-off. After cut-off, pitch programming controlled the missile’s elevation; command guidance continued in azimuth until dynamic pressure became insufficient for control. When the missile crossed the elevation null axis, combined continuous-wave radar and Doppler information were used to compute the desired impact point. This information was stored in the missile as corrective data along with information on departure of the missile from a vacuum trajectory. Re-entry maneuvers up to a maximum of 2G allowed corrective actions to be taken. When the vacuum trajectory was tangent to a line passing through the target, the missile was programmed on this straight line. The terminal phase of flight, therefore, was essentially independent of variation in flight time.
A3A was canceled as a tactical missile in 1951 and the fabricated missiles were assigned as RV-A-8 test vehicles for A3B (XSSM-A-16). The flight test program for A3A missiles testing the propulsion and guidance systems for A3B concluded in January 1954 with the flight of the seventh missile.
Eight A3A missiles were programmed for the flight tests at White Sands Proving Ground; the first round was destroyed by an explosion and resultant fire during static tests at Malta Test Station on 7 August 1952. The initial flight test on 13 March 1953 went well for 23 seconds; at that time turbopump failure caused an explosion and the missile was destroyed. Round #2 on 13 June recorded a missile velocity of 4,450 feet per second at burnout. Changes in tail section construction were affected in Round #3 to minimize spurious burning as noticed during the flight of the second round. The third A3A flight test was 13 August; the missile reached a burnout velocity of 4,150 feet per second. Round #4 was fired on 5 October 1953; higher-than-expected performance of the turbopump and the rocket motor resulted in a cut-off velocity of 3,540 feet per second as compared to a nominal 3,050fps. The excessive velocity was blamed for roll instability occurring at 161 seconds. Round #5 operated at a reduced cut-off time which brought cut-off velocity nearer to normal at 3,083-feet per second. This missile operated as expected until 75 seconds after launch; failure of the 150-volt and 120-volt batteries resulted in loss of flight control from the ground. The tail section of this missile was not recovered, and subsequent firings failed to display similar troubles, so the cause of the battery failure was unable to be determined. Round #6 on 20 November functioned well until about 20,000 feet before impact when a tail fin control malfunctioned, and the missile spun out of its trajectory. Information received on the operation of the radio guidance system in this flight proved valuable, however. The final A3A test flight, conducted on 15 January 1954, also yielded extensive data on the guidance system performance despite the loss of stable flight control 53 seconds after take-off. The second and third rounds, in spite of higher than normal cut-off velocity and loss of ground guidance control early in the flight, operated well until breakup at re-entry at 325 seconds and 315 seconds respectively. Studies performed following the seventh flight were designed to learn to cause of power failures previously noted; power arcs were started and sustained between closely spaced conductors such as the high voltage wiring and water droplets formed on the liquid oxygen pipes. The arcing was cited as possible explanation for the failures.
Previous research and development resulted in the HERMES A3B which was to be the last of the missiles launched at White Sands Proving Ground for the HERMES Project. In physical makeup, this final result of HERMES Project tactical missile development was 400 inches in length, 47 in diameter, had a fin span of 100 inches, and weighed 5,319 pounds without fuel and 11,850 pounds at take-off; its range was 90 nautical miles maximum, although the tested range at WSPG was 61 nautical miles with nominal figures of 119,000 feet in altitude, 3,050 feet per second velocity at cut-off, and 217 seconds of flight time established for the test flights.
Though basically the A3A propulsion system, the A3B power plant incorporated several changes and component improvements. These changes were designed to improve efficiency, to simplify the system, to increase reliability and to cut cost. Propulsion components not actually used in flight were removed from the missile itself and were located on the launch stand whenever possible; the oxidizer tank, fuel tank and peroxide tank operating pressure; and over-pressure switches were so relocated.
The A38 inertial guidance system was independent of any ground-located equipment. It was comprised of a stable platform or odometer system and its associate amplifiers, a range and azimuth unit, and a time-to-go unit. Functions of this inertial system were to:
(1) Accept corrections necessitated by deviations occurring prior to odometer control
(2) Sense accelerations that caused displacement in the impact target coordinates
(3) Provide autopilot channels with displacement signals to correct the trajectory
(4) Provide missile attitude signals
(5) Initiate a time departure from vacuum to terminal trajectory
The radio guidance system was actually two systems: one coarse and one fine. The former was composed of a ground-based tracking radar and a missile-borne radar beacon; the latter included a ground-based S-band transmitter, a Doppler receiver, fine angle receivers, a digital computer and an antenna mount with receiver, decoder, transmitter and oscillator borne in the missile. At WSPG, the ground radio guidance equipment was housed in and about a specially designed building located 9,000 feet south and 1,000 feet west of the launch stand.
Launch and handling equipment included the major items employed in R&D tests at White Sands Proving Ground of the RV-A-8. The launching stand provided support and stability to the missile under winds up to 50 miles an hour; liquid oxygen storage tanks were located about 30 feet from the launching stand and were protected on the missile side by sandbags; the erector was a meilerwagen originally designed and used in the V-2 firings, then redesigned for the A3A series and later adapted to the A38.
Range safety requirements had been discussed for several years prior to the A38 test flights, and positive steps had been taken to insure emergency cutoff facilities in previous test series. Safety developments to the time of the A38 series resulted in a cutoff receiver and antenna mounted on fin #3 of the missile. At receipt of the proper ground-based command, the safety device would trigger a guillotine in the missile; the guillotine then would sever a pneumatic control line to the peroxide valve and also would sever electrical power leads to all four fin servo actuators-the first action terminated thrust and the second left the missile unable to stabilize itself in yaw, pitch and roll.
The flight test program for the A38 was formulated in early 1953 as a series of 28 firings. The first six missiles were to be test-fired in order to perfect an operating system; the second six missiles were destined for tests of accuracy; the remaining missiles would provide data on system performance under simulated field conditions. In mid-1953, the number of A38 missiles authorized for flight tests was reduced to six but the original objectives for the first six missiles were retained to demonstrate the feasibility of the basic guidance system concept.
Flight #1, Round #1 of the A38 was conducted at White Sands Proving Ground on 11 May 1954. The missile air-frame performed in a satisfactory manner until breakup at 192 seconds; propulsion was only four percent under the nominal figure; inertial guidance operated well until 141.5 seconds when a power supply failed; velocity data gathering was terminated early due to misalignment of the ground receiver station antenna. Altitude achieved by this first A38 was 106,000 feet compared with 118,000 feet nominal altitude for this and all succeeding A38 flights.
Although breakup occurred before the nominal impact time of 217 seconds, telemetry signals were received from the mid-tail section until 260 seconds. Major pieces of debris were found in the 48.2 nautical mile area.
The second test flight, Missile Round #3, on 20 July met most expectations. Thrust was slightly higher than nominal and propulsion cut-off was recorded about two seconds earlier than the nominal time of 62.1 seconds. Peak altitude was 117,600 feet, range was 59.7 nautical miles and duration of the flight was 214.6 seconds. The inertial guidance functioned satisfactorily until 165 seconds when loss of roll control was noted; the autopilot drag accelerometer also malfunctioned at this time. A premature zone-entry signal in the ground-based equipment was the only flaw in the radio guidance system.
Round #2 was fired on 26 August; the third in the series. Unintended transmission of erroneous and premature proportional guidance commands resulted in a large deviation from the planned trajectory; the guillotine emergency cut-off device was triggered at 30 seconds and the missile impact occurred eight miles out after 68 seconds.
Round #4 on 21 September operated as expected until about 208 seconds when oscillations of pitch, yaw and roll began because of a malfunction of the drag accelerometer circuit. Proportional guidance commands were terminated early. Missile breakup occurred at 213 seconds. Despite this terminal phase breakup this missile reached an estimated peak altitude of 118,000 feet and impacted, in pieces, in the target area.
Flight #5, Round #5 on 19 October exceeded pre-flight plans. Peak altitude was 118,200 feet, flight range covered 63.3 nautical miles, flight time was 225.2 seconds, and the missile was stable to impact. Inertial guidance and the autopilot system operated properly. Spurious operation of the missile-borne stepping switch was noted in the radio guidance system as was a boresight error in the coarse guidance radar antenna; radio guidance performed satisfactorily otherwise.
The sixth and final A3B was fired on 16 November 1954. Missile breakup occurred at 204 seconds. The inertial guidance system’s azimuth-range acceleration servo opened at 87 seconds and remained inoperative for almost 100 seconds; other components of the inertial guidance system were noted as performing properly. The autopilot system functioned as planned. The radio guidance system was satisfactorily although range data was erratic because of variations in the beacon integration frequency. Peak altitude was 113,000 feet, range 54.5 miles and flight time 206.6 seconds.
Although system accuracy was not demonstrated and, therefore, none of the A38 flights considered fully successful, the series did prove the system inherently capable of the predicted performance based on the actual hardware employed.
HERMES Project technical work not previously terminated was ended on 31 December 1954. The Industrial Division, Redstone Arsenal, supervised disposition of residual materials and facilities. The modified AN/MPQ-12 tracking radar used in the R&D flight tests was permanently placed at White Sands Proving Ground.
Although no formal training program was undertaken in connection with HERMES Project, several efforts were made to utilize HERMES A1 flight tests and, later, HERMES A38 operational missile flight tests at White Sands Proving Ground for training of Army Field Forces personnel.
Verbal information from a HERMES employee indicated that approximately 20 Army officers and enlisted men were assigned to HERMES for on-the-job training at various times. These were technical developments in the project. The technical experience gained by these trainees was basic to guided missiles rather than on specific missile systems.
In December 1951, an A38 Training Conference was conducted to plan for the provision of User and Ordnance personnel to work with the first operational A38 missiles. In early 1952, the contractor submitted a proposal for the training aids, lesson plans and an instruction manual. The HERMES Project status indicated that such training would be premature and no further action was taken in regard to training.