Msfc Program Evolution 1
Figure 11. Project evolution.
- Figure 13. Jupiter configuration.
2. Evolution/Creativity/Innovation. During the Jupiter and Redstone programs, there was a struggle/trade of how best to deal with aerodynamically unstable missile systems, including performance loss due to winds, guidance, and loads. The aerodynamic stability question could be solved by putting large static aerodynamic fins on the aft end, and optimizing the external shapes and mass distributions (propellant tank locations). In many cases, the fins were too large, thus thrust vectoring was brought into play along with movable fins. In the case of the Redstone, movable vanes were put into the thrust stream. The Jupiter vehicle gimbaled the nozzle, driving the design of the gimbaling thrust bearing and actuation system. There were still the performance, guidance, control, and loads problems associate with the winds/aerodynamics-induced drift and loads.
Werner K. Dahm, whose experience spans the German V-2 through the present, provides his assessment of this area of development in the following sections.
C ontrol Development of Rocket Vehicles:
a. During the time of the V-2 (World War II in Germany) and Redstone (figs. 12a and 12b) (after World War II in the United States), the control technology could neither cope with aerodynamic-ally unstable vehicles, nor with aerodynamically unstable control surfaces. The vehicle had to be stabilized with fins of sufficient size to keep the center of pressure behind the center of gravity at all times, and the center of pressure of control surfaces had to be downstream of the hinge line. Primary control was exerted by jet vanes (aerodynamic controls require dynamic pressure to be effective, and are thus useless near lift-off and cut-off). Jet vane size had to be minimized to reduce the thrust loss from their drag. In addition, jet vanes are close to the single-engine-vehicle's longitudinal axis and provide only weak roll control. The effect of side winds on the fins creates "rolling-moments-due-to-combined-pitch-and-yaw" proportional to the dynamic pressure. The jet vanes could not overcome these moments. Therefore, aerodynamic control vanes were added at the fin tips of both missiles. These vanes were coupled with the jet vanes, and were driven by the same actuators. The Redstone had about the same range and payload weight as the V-2. However, the payload was a nuclear bomb, and the circular probable impact error was reduced to 450 m through terminal guidance of the separated entry body, i.e., the biconical nose of the missile. The circular probable error of the V-2 was about 10 times as large. The terminal guidance required:
(1) Attitude control during the exoatmospheric flight of the separated warhead, to keep the control gyros from hitting their stops and losing their alignment. This control was accomplished with cold gas control jets located at the roots of the warhead air vanes. The jet controls were hard-coupled with the air vanes.
(2) An aerodynamic design of the warhead and its air vanes for minimum center of pressure shift with Mach number, in order to minimize control system size and weight. This had to be accomplished on paper, and methods for this had to be developed along the way, since supersonic wind tunnels were not operational during the initial years of the Redstone design. The first warhead test data became available about 4 months before the first Redstone launch (August 20, 1953). The paper design was successful; no design changes were subsequently required.
The Redstone was really two missiles. The initial version, dubbed the "Experimental Redstone," was designed for a large, spherical payload; its design started in early 1951. In 1953, that payload was replaced by a more slender, cylindrical shape, which forced a redesign of the warhead and led to the "Tactical Redstone." That vehicle was eventually deployed in the field.
b. Hermes II: First Attempt to Fly an Unstable Launch Vehicle. The Hermes II project of the late 1940's was supposed to test fly a wing ramjet vehicle at Mach 2.7 and an altitude of about 30 km. The vehicle was a modified V-2. The tip up to the aft end of the instrument compartment was converted to a second stage, with a large wing ramjet, a pitch control canard, and two vertical control vanes at the rear end for lateral and roll control. In spite of the enlarged tail fins, the large ramjet wing rendered the launch configuration unstable. An accelerometer control was supposed to stabilize the vehicle. That control system worked well in bench tests, but failed in flight because the mechanical vibrations of the vehicle structure swarmed the accelerometer. The vehicle lost control somewhat into the flight. The launch took place at White Sands. Due to the severe limitations of the telemetry of that time, the flight vibration environment was largely unknown. The Hermes II project was canceled in early 1950.
c. Jupiter, IRBM, etc.: Swivel Engine Control of Unstable Vehicles. By about 1955, when the design of the Jupiter and Thor IRBM's and the Atlas and Titan Intercontinental Ballistic Missiles (ICBM's) began (all having the common Atlas developed engine), the propulsion community had managed to design swivel engines for thrust vector control. This powerful control authority, together with advances in control technology, permitted us to design and fly aerodynamically unstable missile configurations. Fins could by and large be abandoned. Jupiter became a simple blunt cone-cylinder body; the forward part of the cone was the reentry body, with the rear part being the instrument compartment. The control system included four cylindrical angle-of-attack meters sticking out from the surface of the instrument compartment, two each in the pitch and yaw planes. They compensated for the instability of the basic configuration. The system was successful and reliable. Jupiter roll control was accomplished by using the thrust of the engine's turbine exhaust, discharged through a swiveling pipe knee at the flank of the engine compartment.
The first breakthrough was the development of a drift minimum control logic that balanced the system between control, guidance, and performance in terms of a rigid-body system. This did not solve the induced loads problem. A rigid-body load relief scheme was developed that required schemes to sense the induced angle of attack. During these programs, two approaches were used. (1) Small vanes on the vehicle nose that turned to follow the aerodynamically induced flow were used, with transducers to measure the angular deflection. (2) Delta pressure sensors were placed on each axis of each side of the vehicle to provide a means of detecting angle of attack. In order to really work these issues, a good understanding of the atmospheric winds and density was required. The Air Force and the Army started programs to measure these atmospheric characteristics, requiring development of sensing systems, data evaluation, and modeling technologies.
Liquid propellant sloshing became a problem due to the loss of the second Jupiter vehicle. Analytical and experimental data and approaches were not available to characterize the problem. The first attempt was to place the full-size propellant tank on a railroad car then bump it against the rail end stop as an excitation mechanism. In terms of today's standards, this was not a well-controlled experiment. Large perforated cylinders with sealed spheres (commode floats) were floated in the tank and again tested, showing their good damping characteristics. The next Jupiter was flown successfully using this system. New technologies resulted as the program moved forward:
1. Analytical representation of fluid dynamic characteristics
2. Equivalent mechanical analog of the complex analytical equations (pendulum or mass spring slosh model)
3. Scale model testing techniques, instrumentation, and data evaluation
4. Development of slosh baffles that became an integral part of the structural stiffening ring, saving weight over the floating (beer) cans.
Because the warheads (payloads) had to reenter the Earth's atmosphere at very high speeds, aerodynamic heating and protection against it were major technology developments. During this time, the ablative protection attached to the structural metal was developed, verified, and used, and serves as a base technology for many future systems.
Plume heating of the vehicle base, both convection and radiation, was an additional technology issue. The first Jupiter was destroyed due to plume heating that destroyed the control wires, thus causing it to lose attitude control. Two technologies had to be developed and verified: (1) environment prediction, and (2) protection/control of environment.
Structural dynamic testing of the full Jupiter vehicle was attempted, but was not very successful due to improper boundary conditions (suspension systems); however, this served as a solid foundation for testing technology developed for Saturn.
During the Jupiter development, the interaction between the thrust vector control (TVC) servos and the second bending mode was a major effort. The potential problem was brought home in force on the first flight and subsequently fixed.
According to Harold Scofield, "Some people were never convinced that the second Jupiter lost was due to sloshing. We had only primitive models before the loss in flight. It is difficult to make definitive conclusions without some good analytical models. This problem (lack of definitive models) is still with us in some areas today, thus, exists the dependency on test data alone for those areas where modeling is inaccurate. The lessons keep repeating. We must pay attention to history or repeat it. Theories and analysis are always needed whereby to interpret the test and flight data."
All of this was done with very little computational capability, which points out the innovation and practicality of the engineers in dealing with complex technologies. Those were the days of slide rules and Marchant and Frieden desk calculators.
As was mentioned previously, graphite jet vanes were used in the combustion gas stream (thrust stream) to control the thrust stream and thus the vehicle. However, severe erosion limited its effective control duration. This led to gimbaling the thrust chamber. Gimbaling the thrust chamber with the high-pressure ducts from the pumps input very large actuator loads and short duct life. This led to gimbaling the entire engine with the pumps attached to the engine. Balanced bellow ducts were incorporated for longer life, lower actuator power, and larger gimbal angles.
The Redstone thrust chamber was cooled through the double-shelled construction of 1/8- and 1/4-inch plates and contained several expansion joints. The shell plates were replaced by very long, thin tubes that significantly reduced the engine weight, eliminated thermal buckling problems, and allowed high-rate gimbaling at reduced actuator loads and power. Alcohol and hydrogen peroxide fuel systems gave way to a common kerosene lox propellant system for propulsion and pump turbine power.
It should be remembered that, initially, the Redstone was to be essentially a German V-2 built to American standards by the Army with the Germans brought to this country after World War II. The engine started out as an Air Force development program with a three-phase initiative: to understand liquid propulsion, derive the physics formulation in all associated disciplines, and apply the derived physics of a V-2 type engine using American standards and advanced physics. The final results were 50-percent less weight and 50-percent more thrust. This work was transferred to the Army team, which developed the Redstone. In the end, due to the innovations applied, it evolved to a true defense missile with a large warhead.
3. Problem Examples .
a. Redstone Potentiometer Feedback. The first dynamic problem experienced occurred early in the Redstone rocket program. A Redstone vehicle was in checkout and verification in a horizontal position in its transportation cradle. These early vehicles were manufactured at the Redstone Arsenal, checked out, then transported to Florida for launch. In this case, the control system was activated for checkout. The control sensors had a potentiometer pickup; due to some light shock, the wiper arm was moved from one wire to another, which resulted in a control signal. As a result, the jet vanes moved, exciting a structural mode. The structural mode in turn caused the wiper arm to move back, creating a new signal. The result was a closed-loop limit cycle instability between the sensor (pickup) jet vane (inertia), and structural mode, ringing out at the first mode frequency. The noise of this closed-loop resonance was very loud, vividly demonstrating closed-loop instabilities. The fix was simple in that a filter was incorporated in the loop that filtered out the frequencies associated with the modes and sensor pickup, breaking the loop and stabilizing the system. A later design also changed the pickup to a continuous magnetic type, adding margin to the problem solution.
b. Jimspheres (Atmosphere Sounding Balloons Erratic Response) . Atmosphere environments are key to predicting space vehicle response during the ascent phase. Atmospheric winds are the key parameter to loads, flight mechanics, and control prediction. As a means of developing a statistical quantification of these winds, a balloon radar tracking system was developed by MSFC's atmospheric group under the leadership of Dr. William Vaughan. The goal was to not only measure large-scale or mean environments, but to get an accurate quantification of the wind gust down to 25-m wavelengths. The attempt to measure these small gust effects met with frustration. The smooth skin balloons were unstable (type of flutter or vortex shedding). In a controlled, no disturbance environment, a rising sphere would oscillate (fig. 14). A classical problem most would say. Dr. Jim Scoggins found the solution by observing the golf ball, then instead of small dimples, he added many conical spikes to the sphere's skin. The problem was solved. The sphere was stable (fig. 15). The resulting data base used throughout NASA is evidence. As a result, this modified sphere was named Jimsphere (after Jim Scoggin) as well as the data base (fig. 16). Many people have a small tie clasp with a miniature Jimsphere as a reminder of the agony one sometimes must go through to achieve the innovation required for solution to unexpected problems.
Time lapse trace of Jimsphere ballon released at 11:54 p.m., August 2, 1963 during stable atmospheric conditions and light winds.
Figure 14. Unstable Jimsphere time history.
Time lapse trace of rose ballon release at 11:25 p.m., August 2, 1963, during stable atmospheric conditions and light winds.
Figure 15. Stable Jimsphere time history.
- Figure 16. The Jimsphere balloon wind sensor.
c. Jupiter Sloshing Instabilities. A closed-loop control instability occurred on the early Jupiter firings. The Jupiter was a liquid propelled military vehicle. Sloshing propellant coupled through the control system and became unstable. This instability saturated the control system, and the vehicle went out of control during the maximum dynamic pressure regime of launch and broke up. The results were dynamic, with beautiful fireworks high in the sky, but very costly to the program. The instability was aggravated by the trajectory tilt program. The tilt program was a series of discrete steps instead of a continuous functional change that started the oscillation and reinforced the amplitudes of the wave through a forced oscillation. At this early phase in the rockets and space age, models did not exist for analyzing problems of this type. As a result, several things happened. Propellant sloshing data had to be obtained quickly. No analytical solutions were readily available. A test program was started that included both scale-model and full-scale testing. A slosh suppresser had to be found before the next launch. In order to meet this goal, a full-size propellant tank filled with water was placed on a railroad car. The railroad car was bumped against the track end stop as an excitation source. The first test was without suppression devices. Water was used to simulate propellant to establish frequencies, etc. Various devices were tried next as suppressers. The one chosen was called beer cans, which consisted of long perforated cylinders with flotation spheres at the top. The entire surface of the propellant was covered with these devices (fig. 17). The test showed more than adequate suppression was achieved, and the next launch was slosh free. In the meantime, other solutions were pursued, including development of analytical characterization of the sloshing propellant. This resulted in the development of slosh baffles (rings inside the tank that became part of the structural stiffness (fig. 18)), as the most effective analytical means of suppressing slosh, and parametric test data were acquired for oscillating propellants in both zero- and high-gravity (g) fields. Today, all space vehicles are analyzed and designed with this phenomenon in focus as a potential problem. The lack of analytical and experimental data prior to launching, as well as lack of experience in these type problems, led to the failure of the Jupiter missile due to propellant sloshing control system coupling. The fix was fairly easy and did not impact flight schedules drastically. This is not always the case, and repeats of this type instability should be avoided if possible. The innovative way special tests were conceived and conducted to meet launch schedules should be a lesson in this age of precise testing.
- Figure 17. Beer cans, Jupiter anti-slosh devices.
Figure 18. Slosh baffles, Jupiter missile.
Saturn V SW-IC Fuel Tank Assembly
Figure 18. Slosh baffles, Jupiter missile.
B. Saturn/Saturn Apollo/Skylab Launch Vehicles
The Saturn family was a unique experience in the evolution of space technology.23 Being a family, it provided a building-block approach that led to the successful Moon landing and the Skylab space station program. It was all built on the foundation of the successful Redstone and Jupiter vehicles. The Skylab space program had two distinct parts: (1) the launch systems used to launch both the station the manned visits used modified Saturn V and Saturn IB vehicles, and (2) the operations of the Skylab space station and Apollo telescope mount (ATM). The Skylab launch vehicle is put in this section because of the commonalty of the launch systems. The Saturn family is shown in figure 19a.
1. Characteristics. The Saturn I was envisioned as both a test-bed and a launch vehicle built using available technology and manufacturing tooling. The first stage was derived using a cluster of eight Redstone size and one center tank of Jupiter size propellant tanks and a cluster of eight of the Jupiter engine system. Figure 19b gives its characteristics. The lower ends of the tanks were attached rigidly to a large thrust frame. The engines were attached to the other side of this frame. The upper ends of the tanks were attached using a spider beam that also attached the second stage SIV. The tanks containing the cryo propellants had to have longitudinal slippage to account for thermal contraction, thus the load paths from the first stage engines were through the center tank and the four fixed fuel tanks. The sliding joint on the lox tanks were at the upper end as a connection to the spider allowed for the thermal expansion.
The second stage, SIV, was a liquid hydrogen/lox stage powered by a cluster of RL-10 engines. A very interesting story occurred relative to the first launch of the Saturn I vehicle. The SI was dynamically tested full scale to validate its complex dynamic characteristics. In order to simulate the flight conditions of free-free, the vehicle was suspended on bungy cords (elastic supports) where the vehicle suspended frequency was well separated from its first free-free frequency. In attempting to remove any residual effects of this suspension system on the free-free modes, an analytical technique (filtering of data) was attempted. Due to round-off errors in the data, instead of removing any unwanted effects, the process introduced three false modes in addition to the actual modes. Incorporating these modes into the control stability analysis indicated an unstable vehicle. The initial launch was held up for 2 weeks to sort out the problem. This problem got played in a Fortune magazine feature article. During the 2 weeks, it was possible to show that the modes were indeed false. In a big meeting with Von Braun, the decision was made to launch. The launch was totally successful. Testing approaches and data evaluation are a key element to the success of any program, as was so strongly illustrated here.
The Saturn IB was a derivative of the Saturn I with improvements to the structural and propulsion systems, including performance upgrade (fig. 20). The second stage SIVB had a new single liquid hydrogen/lox engine replacing the cluster of RL-10's in order to get more performance and higher reliability. The Saturn IB was used as the technology and operations demonstrator for the Saturn V Apollo. In fact, they both used the same SIVB stage and essentially the same guidance and control system. The Saturn IB became the launch vehicle for the manned launches to the Skylab space station.
There were two versions of the Saturn V: (1) the vehicle that launched all Apollo Moon missions and (2) the version that launched the Skylab space station. Saturn V Apollo was a three-stage propulsion vehicle plus the manned command and service module and the lunar excursion module (LEM). On later flights, it also contained the lunar rover stowed in the LEM. Figure 21 shows the overall configuration and its characteristics, while figures 22 through 24 are the configuration and characteristics of the three propulsion stages. In the case of the Saturn V Skylab, the SIVB stages were replaced with Skylab space station with only the first two stages being propulsive. This changed the external geometric configuration and thus its ascent loads. Figure 25 shows this configuration and its characteristics. The Skylab itself contained the living features with life support for the crew, a docking module, the ATM, and solar arrays on both the ATM and the workshop for power to operate all the systems. Skylab was a very successful program. After the four science missions were over, it
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