During FY 1971 the direct responsibility for Space Biotechnology programs was transferred to a new centralized Biotechnology office with OMSF. The shuttle biotechnology effort is incorporated within the overall shuttle technology program by the Working Group and the Shuttle Technologies Office but OAKT R&D funds are not involved during FY 1972.

SUMMARY

During FY 1972 OART will continue its strong technological program to support the Agency's space shuttle program. It does so through the OART Shuttle Technologies Office and its highly effective ad hoc technical Working Group organization by assembling and directing expert disciplinary talent at the Research and Space Flight Centers. OART devotes its expert personnel, its unique faciilties, and appropriate portions of its funding resources to this focused technology program.

The program plan is generated by technological experts within NASA Centers and honed by a series of reviews involving OART and OMSF headquarters offices and Center managements to maximize results that can be achieved with available resources. Since every review along this complex line involves NASA's most experienced research and technology people, the result is a highly refined, concentrated, hard-hitting, and timely series of efforts that will support successful development of a reusable space launch vehicle.

Aerothermodynamics work has defined and selected the characteristic shuttle configurations. Major concentration has been put on structures, thermal protection systems and materials because of the key importance to the airframe of shuttle weight and cost. The area of vehicle dynamics and aeroelasticity is being augmented progressively as the shuttle becomes increasingly well defined. In propulsion special emphasis has been placed on the hydrogen-oxygen auxiliary propulsion. While it appears that no major technology deficiencies in electronics stand in the way of system development, our principal intent in this area is to define the entire integrated system in order to reduce development and operational costs. Somewhat smaller technology programs are being pursued in the area of operations, maintenance, and safety where the principal emphasis is on the handling and management of cryogenic liquids, and in biotechnology, which is concerned with the well-being of the crew and the functional integration of machine and crew in the shuttle operations.

Substantial progress has been made in implementing this well organized and coordinated shuttle technology program. The program has been planned to meet first priority needs on time. Certain initial phases of the program, notably aerothermodynamics, leading to configuration selection, advanced superalloy development and testing, and definition of a preferred auxiliary propulsion system type have been substantially accomplished on the planned schedule. Other elements are progressing.

The overall planned technology program will have been fully implemented during the second half of FY 1971. The technical results are disseminated to the aerospace industry. USAF, and all involved elements of NASA, by a series of contractor and NASA reports and by conferences. A series of four technical conferences, each in special technical areas, will be held during the spring months of 1971. The first of these, on aerodynamics, structures and materials, and vehicle dynamics and aeroelasticity will be held at the Langley Research Center on March 2, 3, and 4.

PREPARED STATEMENT OF WILLIAM H. WOODWARD, DIRECTOR, SPACE PROPULSION AND POWER, OFFICE OF ADVANCED RESEARCH AND TECHNOLOGY, NATIONAL AERONAUTICS AND SPACE ADMIN

ISTRATION

INTRODUCTION

Mr. Chairman and Members of the Committee:

The general objective of the Space Propulsion and Power Program is to provide the technology base necessary for the ultimate development of systeins to meet NASA's future mission requirements. Emphasis in FY 72 in the propulsion area will be on technologies for planetary exploration and auxiliary propulsion for the shuttle with a modest effort aimed at new concepts. Emphasis in the power area, which encompasses solar and chemical technologies and power processing, will be on programs to obtain major increases in performance (2 to 10 times current levels) and reductions in system cost.

Last year, NASA placed all of the solar and chemical propulsion and power programs under the direction of the Space Propulsion and Power Division at Headquarters. The majority of the programs will be implemented by the Lewis Research Center, Jet Propulsion Laboratory, Langley Research Center, and Goddard Space Flight Center, with small but critically important efforts at the Manned Spacecraft Center and Marshall Space Flight Center mostly in shuttle related technologies. Coordination of these programs with related Department of Defense efforts wi'l continue to be carried out through the JANNAF Chemical Propulsion Group, the Aeronautics and Astronautics Coordination Board, and the Interagency Advanced Power Group.

LIQUID ROCKET PROPULSION

Almost every man-made object placed in space-from the first Explorer to Apollo 14 and from Sputnik to Luna-has employed some liquid rocket propulsion. The genius of Dr. Goddard initiated a technology that first made the term "Space Age" a possibility and then a reality. But the technology of Saturn, of Centaur, and of Lunar Excursion Module cannot satisfy the demands of the space shuttle, of outer planet orbiters, of space-based reusable tugs. Very real and very demanding problems continue to confront liquid rocket propulsion. The solutions to these problems are, in fact, the primary goals of this portion of the Space Propulsion and Power Program.

The primary objectives of liquid rocket research and technology activity for FY 1972 are: (1) to provide propulsion technology that maximizes performance for multi-year duration planetary missions; (2) to provide technology for predictable and efficient transfer of cryogenic propellants from one spacecraft to another; (3) to provide an understanding of certain rocket phenomena (e.g., combustion instability) that can preclude problems by design thereby significantly reducing the costly cut and try process currently required; and (4) to continuously evaluate the profitability of new concepts and applications of new materials and/or manufacturing processes to liquid rocketry.

HIGH PERFORMANCE SPACE STORABLE LIQUID ROCKET PROPULSION Future planetary missions impose unique requirements on the spacecraft propulsion systems, principally due to the long transit times associated with interplanetary travel. Reliable and predictable operation of both primary and auxiliary propulsion systems over periods of up to 7-10 years at long distances from earth is required. The functions of the spacecraft propulsion systems during interplanetary travel include: (1) spacecraft attitude control for such purposes as communications, guidance and navigation, propellant thermal control, and scientific observations; (2) trajectory corrections for maintaining the proper flight path to the target planet; and (3) spacecraft maneuvers in the vicinity of the planet to position the payload in the desired planetary orbit and/or on the planet's surface, and in some cases, even return it to earth.

The attitude control and trajectory correction functions (1 and 2 above) are typified by low thrust, low total impulse requirements. Decomposing hydrazine used as a monopropellant appears to best satisfy these requirements. Such a system has been designed to meet the requirements as specified for the JPL. Thermoelectric Outer Planet Spacecraft technology project (TOPS). The principal technology advance has been the design and operation of attitude control hydrazine thrusters in the .050-.100 lbs. thrust range (Figure 1). The high ratio of thruster mass to propellant flow rate extends beyond the current knowledge of the catalytic decomposition process in both transient and steady state operation. Principal effort extending into FY 72 will be to characterize this process analytically for single thrusters using actual hardware design and empirical test data as a basis, and then demonstrate system operation of a cluster of thrusters under simulated space conditions. Although the TOPS program has provided the focal point for establishing realistic mission requirements, this monopropellant hydrazine technology will be equally applicable to other planetary missions, e.g., Mercury, Venus, Mars, and to long-lived earth satellite spacecraft as well.

For propulsion maneuvers in the vicinity of the planet, either in planetary orbit or to and from the surface, the spacecraft propulsion system choice depends on a number of interrelated, critical parameters such as the desired scientific information and associated instrumentation, communications and power requirements, the launch vehicle capability, the target planet, the planetary orbit of interest and the environment in which the system must survive and operate reliably.

The larger, higher thrust spacecraft propulsion systems provide maximum payload capability when pump fed engines are utilized. Mission analyses indicate a number of potentially attractive propellant combinations are possible. The FLOX/CH, (Fluorinated Oxygen/Methane) propellant combination has proven very adaptable to the planetary missions due to its excellent space storability and suitability to pump fed engine operation. The liquid hydrogen/oxygen or liquid hydrogen/fluorine propellant are attractive, provided long-term hydrogen storability is proven feasible. F2/H2 (fluorine-hydrogen) would maximize payloads while oxygen-hydrogen would circumvent the problem of handling the highly energetic fluorine propellant. Each of these systems is also readily adapt

able as an upper stage for greatly increasing the payload delivery capability of the existing stable of NASA launch vehicles, or as earth escape stage carried into earth orbit by the space shuttle.

The current program at the Lewis Research Center is directed towards two technology areas common to whatever propellant combination is finally selected: (1) storing propellants for long periods of time in the spacecraft operational environment; and (2) pump fed engines for spacecraft and upper stage applications for a typical cryogenic bipropellant stage design (Figure 2). In FY 72 the storability of liquid hydrogen will be evaluated employing the most advanced techniques in super insulation, minimum heat-leak tank supports and penetrations and shadow shields. Liquid nitrogen will be used as a simulant for the oxidizer in conjunction with the liquid hydrogen. Tests will be conducted in a vacuum chamber simulating the space environment. During FY 71 all components of the system were procured, and subsystems assembly initiated. Subsystem testing will be initiated in FY 12 leading to complete system testing early in FY 73. In the pump fed engine area, a contract was awarded in FY 71 for the design, fabrication and test of a 5000 lb. thrust, FLOX-Methane engine, based on prior successful component evaluations and engine design studies (Figure 3). The space storable FLOX-Methane was selected as the propellant combination since this technology area is directed primarily towards advanced, large planetary spacecraft anticipated for the more distant future. Subsystem and system tests are planned for the FLOX-Methane system as a typical space storable combination, following the aforementioned hydrogen storability tests in FY 72 and 73. A direct comparison then will be made on propellant thermal control techniques between the deep cryogenic and the space storable propellants. The 5000 lb. thrust engine will be added to the propellant system and tests conducted simulating an entire mission, from launch to planetary orbit insertion including engine firing. Again, although the propellant choice is FLOXMethane, the system engineering information gained will be applicable to a significant degree to any advanced pump fed engine propulsion system, whether for planetary missions, upper stage applications, or a reusable space tug.

For some of the more difficult near-term planetary missions, launch vehicle limitations dictate relatively small spacecraft. Thrust levels range from 500 to 1000 pounds thrust and propellant quantities range up to 3000 pounds. In this weight and thrust class of propulsion systems, pressure fed engine operation is optimum for maximizing payload. In the past, the earth storable N2O4/MMH (Nitrogen Textroxide/Monomethylhydrazine) propellant combination has been used in these types of missions and will be used in the Mariner 71 Mars Orbiter mission and Viking 75. For the more difficult outer planet orbiter missions to Jupiter and Saturn, higher energy propellants are attractive for maintaining adequate payload capability.

In addition to providing high performance in pressure fed engines, candidate propellant combinations must also exhibit good space storability in small quantities. Three propellant combinations have emerged as most applicable to this class of missions. The OF2/B2H. (Oxygen Diflouride/Diborane) combination offers the highest performance potential, while FLOX/MMH (Fluorine Oxygen-Monomethylhydrazine) would provide the advantage of the least modifications to the operational N2O4/MMH systems. The third combination of interest is F2/N2H4 (Fluorine/Hydrazine). Here the hydrazine could be used for both the primary orbit insertion bipropellant propulsion system as well as for the attitude control and trajectory correction monopropellant systems. Thus the orbit insertion engine and the trajectory correction engine would be combined into a single unit with dual mode operating capability eliminating one propellant tank and one thruster plus associated valving and plumbing.

Current technology efforts in the pressure-fed area are directed towards longterm thermal control of small propellant quantities and the acquisition of information that will permit the design of durable, high performing flight configuration engines. The ultimate objective of this work is to demonstrate the operational capability of an integrated spacecraft propulsion system, over a complete simulated mission, from launch to planetary insertion. The FY 72 program will emphasize engine design evaluations for the purpose of selecting a single propellant combination for initial integrated system testing. Representative propellant tank sizing, thermal control and feed system components are being designed and evaluated to accommodate any of the candidate propellant combinations in order to minimize cost and to maintain the projected technology schedule. This effort is directed towards providing technology readiness in time for planetary orbiter missions commencing in the late 1970's and continuing through the 1980's and beyond.