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U. S. Army fixed wing aircraft employed in R. V. N. at the present time range in sophistication from the 0–1, “Bird Dog”, a single engine, two place light aircraft, to the OV-1, “Mohawk”, twin engine, two place turbo-prop aircraft. Each aircraft has unique escape problems and employs different methods that are appropriate to meet the requirement. U. S. Army rotary wing aircraft in R. V. N. range in sophistication from the OH-13, “Sioux”, to the CH-47A, “Chinook”. All rotary wing aircraft present a formidable escape system design problem. Attempts to design such a system have been made, but none have become available for operational use. Therefore, no single escape system operation at the present time is appropriate for all Army aircraft. Current practice prohibits the employment of an escape system for crew members that is not available for passengers in Army aircraft, for example, the aeromedical evacuation aircraft.

The use of hydraulic tubing permanent joints, when using the induction brazing method, has proven feasible in the manufacture of aircraft. The use of the permanent-joint method in place of mechanical joints has resulted in a reduction of component weights and a lessening of maintenance problems associated with hydraulic systems. The reasons behind the decision to use permanent joints in the A-7A aircraft and the search for suitable methods and materials is outlined. The testing and application procedures of the permanent-joint system and the results obtained to date are given in detail. The purpose of this paper is to outline the approach used by LTV Aerospace Corporation for the adoption of hydraulic tubing permanent joints, when using the induction brazing method, for installation on the A-7A Corsair II; the first full-production-model aircraft to incorporate such a system.

In the design of fluid systems plumbing for aircraft engines, the designer is continually challenged by the problem of component location and routing. In order to achieve accessibility and maintainability, and to avoid physical interferences, plumbing design is accomplished through extensive graphical projection and mockups. The solution to this problem must also satisfy tubing stress limits, resonant frequencies, bracket or clamping positions available and future space requirements. To facilitate and expedite this design procedure a digital computer technique has been developed which determines the clearances between tubing and other engine components. Though not a substitute for graphical projection, this program provides a means for accurate checking for interference. It also serves a valuable purpose in the storage of previous or alternate plumbing routing arrangements for comparison.

A method is described for estimating the influence of thermal gradients on separable tube connectors. Analytical descriptions are given of the connector structural elements which sustain tension, compression, bending, and shear. Using thermal gradient data, and the analytical description of structural elements, a method is presented for including the effect of thermal gradients in the design of high-performance tube connectors. An example is given of the design method applied to a conventional bolted flanged-type connector. The method is not limited, however, to this connector configuration. Correlation of calculated effects of thermal gradients with measured results is presented to verify the design method.

This is a state-of-the-art report on weldable tube fittings, welding equipment and inspection methods for connecting tubing in aerospace aircraft hydraulic systems. Design advantages and limitations of several types of fittings, including a new swaged-on concept, are discussed. Initial flexure-endurance tests of the new design indicate greater than a 30-fold improvement. Tungsten inert gas (GTA) welding equipment utilizing an orbital type welding head with automatically programmed time, current, and gear speed provides remote in-place welding capability. Nondestructive radiographic and ultrasonic inspection methods are also described.

The development of advanced manned lunar systems requires increasing simulation fidelity to verify man/machine interfaces. Lockheed has been involved in the conceptual development of a mobile lunar laboratory and a small cabinless vehicle designed to transport an astronaut plus scientific equipment cargo on the moon. A lunar environment simulation test bed has been developed and is being utilized extensively to establish man's performance capabilities and limitations on the lunar surface. This simulation test bed includes a capability for testing the interactive effects of one-sixth gravity, lunar surface texture, the reduced pressure of vacuum and a treadmill for simulating extended walking tasks. Initial data obtained with the partial gravity simulator are presented.

Simulated hydrospace is a valuable tool for the solution of engineering problems in the area of ocean engineering. It is employed not only in the research into the behavior of materials and structures under deep ocean environment but also in the quality control of industrial products for hydrospace market. Since a simulated hydrospace facility represents a considerable investment, it is important for all concerned in the decision making process for the acquiring of such a facility to understand the function that such a facility performs.

A description of the largest Acoustic and Vibration Test Facilities in existence for the simulation of major launch vehicle dynamic environment is given and the operational characteristics of both are discussed. Sinusoidal and random excitation techniques are described and unique vibration control methods presented. A comparison of the effects of vibration and acoustic excitation on major space vehicle structures is made.

In order to prove the design validity of SST flight control and hydraulics systems prior to flight, the systems will be assembled and operated in testing fixtures under simulated operating loads and expected flight environments. Clearance for first flight will depend upon successful demonstration in the testing complex. The program is named CODE (from controls DEvelopment), and for the prototype airplane it is divided into two stages. Initially, component problems will be studied in a Preliminary CODE fixture using a triple-channel conceptual servo actuator under mechanical and electrical command. This Preliminary CODE will allow for variation in load inertia, simulated aerodynamic load, and in the compliance used to represent the flexible airframe. Flutter loads will be generated by an external force servomechanism. Data from the Preliminary CODE program will guide design of flight hardware for the prototype airplane.

Surface tension principles applicable to the design of propellant management devices for static retention and dynamic control of space propulsion system propellants are discussed. The characteristics of ideal propellant retention and control (management) systems are introduced as criteria for judging proposed designs. Three typical missions are examined and are functionally characterized as (a) large tank high thrust systems, (b) long-life low-thrust systems, and (c) minimum center-of-gravity travel systems. The leading mission requirements and constraints and their influence on the design of capillary propellant management systems are discussed. Flight and design analysis experience are presented and simulation test techniques introduced.

A need exists for some quantitative measure of an aircraft's support or flotation requirements on the ground. Through use of established relations between airfield or landing strip strength and aircraft ground loading characteristics, aircraft ground flotation requirements have been developed which permit assessment of the flotation characteristics of any aircraft or proposed aircraft design. Comparisons between aircraft or setting of minimum flotation requirements for proposed aircraft are made possible by use of these developments. This paper presents limited background on ground-flotation criteria development and some of the concepts on which current criteria are based. It gives an insight into various problem areas in development of the criteria and explains the extent or limits of applicability of flotation criteria presently in use.

Fluid systems are receiving widespread usage in space vehicles. Typical applications include the use of pneumatic power for attitude control of an orbiting satellite. Selection of a pneumatic system for satellite control is predicated on certain key requirements and constraints imposed upon the system. Chief among these are thrust level, total impulse, electrical power, weight, reliability, and cost. The above parameters are considered in a case example of tradeoff studies that ultimately leads to a pneumatic system design for attitude control of a long-life orbiting satellite.

Difficulties in retaining pressurized gases in Agena separable-connector tubing systems dictated a permanent joint solution. Vehicle considerations indicated brazed joints for application to this particular family of space vehicles. Selection of joining method, application problems, and interesting test results are discussed. Projection of the effects of permanent joints, welded or brazed, on system component design and testing in the immediate and distant future are presented.

Application studies of fluidic technology have demonstrated its adaptability and versatility in several ramjet control systems. Schemes for mechanizing three typical controls with fluidic elements to the extent now feasible are presented. Fluidics are incorporated in sensing, logic, and actuation functions. Digital and analog techniques areemployed. The three controls are: the inlet control which optimizes performance through modulation of the center-body position, the fuel control which maintains the desired fuel-air ratio, and the coolant control which manages the fuel distribution in the engine structure to maintain structural integrity. Performance analyses and test results are included.

A preliminary analytical investigation of a flueric solar attitude control device performing the functions of both sensor and actuator was conducted. The system cycle was based upon only incident solar energy as the power source and heat rejection by radiation to space. The working fluid continuously experiences a change of phase from liquid to vapor and recondensation to liquid. The fluid in the liquid phase is pumped by surface tension forces in a porous wick material within the annular volume and also provides large surface area for the heat exchange function in the evaporator. As a result of angular misalignment from the sun direction an unbalanced mass flow develops. From consideration of the redistribtuion of angular momentum (in the sense of a momentum flywheel) it was shown that the vehicle can be rotated to reestablish alignment with the sun direction. The numerical results indicate the feasibility of the concept for space applications.

The application of space-power technology to a possible progression of manned space flight requirements is presented. Of particular concern is the spectrum of mission requirements and constraints which strongly influence the selection and integration of a space-power concept with manned spacecraft, and the desire to “build-in” as much intrinsic value as possible into hardware decisions. An orderly progression of the manned space flight activity can be postulated which extends from immediate exploitation of Apollo earth-orbital capability to manned exploration of the planets. A multipurpose space station occupies a predominant role within this progression and will provide the major stimuli for the development of new and improved space-power concepts.

Environmental conditions in aerospace systems demand the use of metallic seals and the complete absence of leakage. A unique metal-to-metal seal called the Bobbin seal utilizing the elastic and plastic response characteristics of the seal structure is described. Experimental evaluations with 250 specimens, with 99% sealing helium at leakage rates less than 7 × 10−7 atm cc/sec, indicate that the Bobbin seal can overcome almost any problem of leaking joints in missile, aerospace, or aircraft systems. Results obtained with various materials, in a broad environmental spectrum, with tube sizes up to 16 in. are described.

The problem of establishing the configuration and sizing of the primary structure of a large area solar array yielding 20 w/lb is analyzed. Factors affecting the structural design and examined in this paper include: 1. Effects of launch environment. 2. Deploying and deployed phases of a typical mission. 3. Impact of ground support and test equipment upon the design. 4. Fabrication problems affecting structural design. The analyses consider dynamic and internal loads, temperature distribution and thermal control, selection of materials, weight distribution, and interface of the structural and electrical technologies. It is found that the feasibility of design of a 20 w/lb large area solar array is within the present state-of-the-art only if beryllium is used to transmit primary loads.

Measurements were made of helium permeation rates for O-rings subjected to helium pressure on one side and vacuum on the other. Helium pressures up to 1000 psi were used in the investigation. Gland designs included grooves for O-ring squeezes from 25 to 75 percent, dual O-ring grooves, and two tongue-and-groove configurations. Calculations were made of the expected permeation rates based on material permeability data. The results of this study indicate that dual O-rings and close-fitting tongue-and-groove designs provide permeation rates that are lower than the rates that are calculated by conventional methods.

OV-10A LANDING GEARS - The landing gears of the OV-10A have been designed to meet extremely severe landing and operating conditions. The articulated main gear and semi-articulated nose gear assemblies have proven to be well suited to meeting the requirements for high sink rate landings onto bumps, steps and holes in addition to taxiing and making take-offs over undulating contours. Development data on the landing gear shock struts in laboratory airframe drops and actual flight test landings show good agreement with predicted performance.