SAWE Technical Papers

@inproceedings{3301,

title = {3301. The Harpoon Missile System: 30 Years of Preeminent Naval Defense},

author = {Thomas M. Schultz},

url = {https://www.sawe.org/product/paper-3301},

year  = {2003},

date = {2003-05-01},

booktitle = {62nd Annual Conference, New Haven, Connecticut},

pages = {45},

publisher = {Society of Allied Weight Engineers, Inc.},

address = {New Haven, Connecticut},

abstract = {The date is June 21, 1971, and the headline reads ? ?Navy Selects McDonnell Douglas as Prime Contractor for New Anti-Ship Missile.? 

Thus began the legacy of the Harpoon family of missiles more than 30 years ago. 

Originally dubbed the Air Launched Ship Attack Missile, or ALSAM, Harpoon has evolved into the U.S. Navy?s primary anti-ship missile and the longest-running line of cruise missiles ever produced. 

Unsettling world events in the 1960s ? including the Soviet Union stockpiling missiles in Cuba ? convinced the U.S. government that America must be prepared to counter greater tactical threats. The idea for a common missile for both attack and patrol aircraft was born. 

The U.S. Navy wanted one weapon to satisfy two modes of operation ? air-launched attack on surface vessels in forward areas and surface ship attack against hostile surface ships and patrol craft. Five teams submitted bids in response to the Navy?s request for proposal, and the winner was the McDonnell Douglas team. In December 1972, just 18 months after contract award, the first guided Harpoon, an air?launched missile, hit its target. Through the years Harpoon has proven to be a reliable missile for the U.S. fleet and its allies with a 100-percent success rate in combat. 

Thirty years after that initial contract award, McDonnell Douglas ? now merged with Boeing ? continues to produce the Harpoon missile for sale internationally. The Standoff Land Attack Missile (SLAM), a Harpoon derivative, has responded to threats in both Operation Desert Storm and Bosnia. During Desert Storm, SLAM earned the reputation for being the missile that could ?see? its target with such precision that a second SLAM missile could enter the hole created by the first. 

The SLAM Expanded Response (SLAM ER) is now in production. SLAM ER retains the reliable Harpoon propulsion system, but provides a completely new precision guidance system, new integration with the SLAM seeker and advanced data link, new reactive titanium core warhead, and deployable planar wings. SLAM ER provides the Navy with a reliable, lethal, survivable missile system that is easy to employ outside the reach of enemy air defense systems. 

And now Block II has added the precision strike capabilities of both the SLAM ER and the Joint Direct Attack Munition (JDAM) to the proven Harpoon design. This upgrade allows the Harpoon to be both an anti-ship and land attack dual-mode weapon. 

Affordable, low risk, combat-proven Harpoon and Harpoon derivatives. Thirty years and counting.},

keywords = {14. Weight Engineering - Missile Design},

pubstate = {published},

tppubtype = {inproceedings}

}

@inproceedings{1877,

title = {1877. Weight Control of a Large Space Booster},

author = {G R Matheny},

url = {https://www.sawe.org/product/paper-1877},

year  = {1989},

date = {1989-05-01},

booktitle = {48th Annual Conference, Alexandria, Virginia, May 22-24},

pages = {25},

publisher = {Society of Allied Weight Engineers, Inc.},

address = {Alexandria, Virginia},

abstract = {The Titan III had successfully performed many missions placing satellites in orbit from both coasts. The challenge was to design a longer ''stretched'' vehicle by lengthening the fuel and oxidizer tanks on both stages while holding the diameter constant at ten feet. The new Titan IV vehicle would have to withstand larger P-equivalent loads due to the increased weight required by structural beefup to the skin, stringers, and frames. Different sections of the vehicle were assigned different amounts of beef-up based on their relative axial load-carrying capabilities. An initial design vehicle was established. Upon further analysis, it was determined that the increased vehicle and payload fairing lengths imparted a large bending moment on the booster at design Q-Alpha-Total flight conditions. This caused a vehicle redesign, and it was found that stiffness rather than axial load became the design driver in most areas. As the new stiffness requirements were translated into structural detail sizing and analyzed by the mass properties group, the weight problems began to surface Target weights had been established in the proposal phase and revised due to the stiffness criteria, but both values were being exceeded in design. A method to control vehicle weight was urgently needed. The paper discusses how this challenge was met by analyzing critical frames, stringers, and skins with a target weight versus stress margin comparison. Details are presented to show how items became identified for weight reduction and results are quantified. A significant performance gain resulted from this weight reduction program. The excellent teamwork that produced the weight savings has carried over into subsequent phases of vehicle development. Various design groups were united by this effort, and for a while everyone became a ''weights engineer'' and worked for a common goal.},

keywords = {14. Weight Engineering - Missile Design},

pubstate = {published},

tppubtype = {inproceedings}

}

@inproceedings{1815,

title = {1815. Predicting Radial Center of Gravity Statistical Distributions for Ballistic Reentry Vehicles},

author = {L Edington},

url = {https://www.sawe.org/product/paper-1815},

year  = {1988},

date = {1988-05-01},

booktitle = {47th Annual Conference, Plymouth, Michigan, May 23-25},

pages = {25},

publisher = {Society of Allied Weight Engineers, Inc.},

address = {Plymouth, Michigan},

abstract = {The requirements for increased performance from each new generation of spin-stabilized projectiles has forced the need for more accurate mass properties predictions than ever before. Today the mass properties engineer must not only predict the nominal properties, he must also predict the variations about the nominals caused by manufacturing tolerances and measurement uncertainties. To accomplish this requires both a statistical data base from previous programs and the mathematics to combine and/or modify the data to account for the differences in material, mechanical designs and geometric tolerancing. The mathematics to combine component and subsystem uncertainties into total system uncertainties exists and has been well documented in SAWE technical papers as well as in the general literature. However, mass properties and their uncertainties are usually calculated in Cartesian coordinates while ballistic projectile uncertainties are often needed in polar coordinates. To convert to polar, the pitch and yaw axes center of gravity statistical distributions must be combined to create a radial center of gravity statistical distribution. This paper presents a brief discussion of the flight events of a ballistic projectile entering the earth's atmosphere and of the significance of the various mass properties during these events. The importance of the statistical distribution for radial center of gravity is discussed in detail, and a mathematical method for deriving this distribution from the pitch and yaw axes distributions is presented. A Monte Carlo method is used to demonstrate the validity of this approach. The mathematical derivation for a special case of this radial center of gravity distribution is shown and the application of this special case to influence vehicle design is also discussed.},

keywords = {14. Weight Engineering - Missile Design},

pubstate = {published},

tppubtype = {inproceedings}

}

@inproceedings{1760,

title = {1760. Sizing Missile Guidance Systems},

author = {J Pierson},

url = {https://www.sawe.org/product/paper-1760},

year  = {1987},

date = {1987-05-01},

booktitle = {46th Annual Conference, Seattle, Washington, May 18-20},

pages = {65},

publisher = {Society of Allied Weight Engineers, Inc.},

address = {Seattle, Washington},

abstract = {The development of weight estimation methodology for missiles has been, for the most part, a neglected activity. On the plus side, sound tools have been built for estimating missile structures. However, methods for sizing missile components are typically quite crude and obsolete. This paper presents the results of an effort to develop better, more analytical methods for sizing components used in radar and infrared guided missiles. Included are procedures for estimating the weight of radomes, radar antennae, radar transmitters and receivers, infrared seekers, autopilots, computers and battery requirements. The primary objective was to build weight estimating tools which are analytical in nature, base on engineering relationships important to the design and function of the components. In cases where it was not possible to develop analytical methods, empirical methods were devised using constrained regression analysis techniques. The missile industry has progressed at a rapid pace in the packaging and miniaturization of missile components, especially with respect to electronics. This required the development of technology factors to apply to the empirical data allowing the combining of data spanning multiple technology eras. The methods presented provide a means for sizing future missile guidance components. In addition they provide a view back into the past, dramatically demonstrating the tremendous advancements which have been realized in guidance components. much more work remains to better complete the packaging of missile sizing methodology. The results presented in this paper represent an advancement but are not conceived as complete and mature. More effort should be expended, augmenting and even replacing these when better methods are found. Numerous other missile components such as propulsion, hydraulic and pneumatic power systems, and warheads also need support.},

note = {L. R. 'Mike' Hackney Award},

keywords = {14. Weight Engineering - Missile Design, Mike Hackney Best Paper Award},

pubstate = {published},

tppubtype = {inproceedings}

}

@inproceedings{1241,

title = {1241. Aspects on the Configuration Optimization Process of Stike - RPVS},

author = {E S Klinke and E Rutzen},

url = {https://www.sawe.org/product/paper-1241},

year  = {1978},

date = {1978-05-01},

booktitle = {37th Annual Conference, Munich, West Germany, May 8-10},

pages = {11},

publisher = {Society of Allied Weight Engineers, Inc.},

address = {Munich, West Germany},

abstract = {In order to establish a reasonable Strike-RPV-concept, consideration is to be given to the sizing features such as: 

- Navigational accuracy 

- Weapon delivery accuracy 

- Simplified Structural Design 

An other important design requirement is mission flexibility, which results in a high degree of modularity. 

Results of investigations carried out so far and their effects on mass and cost will be presented. 

The strike mission is the most demanding task for remotely controlled vehicles. The integration of contradicting requirements such as: 

- high speed low level flight to survive 

- precise target acquisitation under bad weather conditions 

- effective weapon delivery with small payloads 

This paper outlines some aspects of how to attack the major problems o f the RPV in the strike role.},

keywords = {14. Weight Engineering - Missile Design},

pubstate = {published},

tppubtype = {inproceedings}

}

@inproceedings{1249,

title = {1249. A Review of Current Microthruster Technology for Synchronous Satellite Applications},

author = {H C Krieg},

url = {https://www.sawe.org/product/paper-1249},

year  = {1978},

date = {1978-05-01},

booktitle = {37th Annual Conference, Munich, West Germany, May 8-10},

pages = {18},

publisher = {Society of Allied Weight Engineers, Inc.},

address = {Munich, West Germany},

abstract = {Soon the 1980's will herald a new era in space exploration. The high note will be the inaugural flight of the Space Shuttle transportation system, which should culminate in the establishment of a permanent orbiting space station. In support of these and other operations a complex network of communications and meteorological satellites must be established. These satellites fall into a unique class in that they function at geosynchronous altitudes to maintain their discrete position over the earth's surface, orient their observation equipment, and hold antenna alignment. High performance propulsion systems, capable of reliably providing precise impulsive control over extended periods, are mandatory.},

keywords = {14. Weight Engineering - Missile Design},

pubstate = {published},

tppubtype = {inproceedings}

}

@inproceedings{1160,

title = {1160. Influence of Deployment Sequence on Propulsion System Design Requirements},

author = {H C Krieg and J G Ferguson},

url = {https://www.sawe.org/product/paper-1160},

year  = {1977},

date = {1977-05-01},

booktitle = {36th Annual Conference, San Diego, California, May 9-12},

pages = {24},

publisher = {Society of Allied Weight Engineers, Inc.},

address = {San Diego, California},

abstract = {A computer program has been developed to determine the real time variation of vehicle mass properties during simulated flight analysis. In addition to mass property calculations, the program has the capability of providing a 'best case' deployment sequence for multiple payload vehicles. 

The effect of mission profile and deployment sequence as related to propulsion system design requirements are reviewed for two sample multiple payload deployment vehicles. For each vehicle studied, the major concern was the placement of 

the main axial engine and its effect on other propulsion system components due to induced gumball angle requirements resulting from the sequential deployment of payload.},

keywords = {14. Weight Engineering - Missile Design},

pubstate = {published},

tppubtype = {inproceedings}

}

@inproceedings{1107,

title = {1107. Deployment Vehicle Mass Properties Model},

author = {H C Krieg and C M Mead},

url = {https://www.sawe.org/product/paper-1107},

year  = {1976},

date = {1976-05-01},

booktitle = {35th Annual Conference, Philadelphia, Pennsylvania, May 24-26},

pages = {30},

publisher = {Society of Allied Weight Engineers, Inc.},

address = {Philadelphia, Pennsylvania},

abstract = {A comprehensive math model of a Deployment Vehicle has been prepared based on a Disc-type configuration. Equations and logic presented may be incorporated into other configurations with little or no modification. The developed model incorporates logic to: 

- Select and size major Propulsion Subsystem Components 

- Physically locate and define the mass and volume requirements for al1 major components (i .e. main and ACS engines , tankage, etc.) within the vehicle. 

- Compute cg locations of a1 1 major components and the overall vehicle's cg, moments and products of inertia. 

- Define the vehicle's shroud eject system. 

- Compute the arrangement of various payload configurations within the geometric envelope of the shroud. 

- Compute the most optimum deployment sequence, based upon overall vehicle cg and product of inertia. 

The program, prepared in standard Fortran language, is designed to operate as a 'slave'' program (closed loop) when coupled t o a flight performance prediction model, or as a design program (open loop) for on-line evaluation of candidate vehicles. When operated i n the open loop, a performance event/time sequence is required. This sequence is derived from data generated by the earlier flight performance prediction runs and is uti1 i zed as a representative sequence. The program is divided into four major subroutines headed by a sequence routine which oversees and directs operation i n accordance with the subject instructions. The sequence also gathers and makes available to the subroutine all necessary data (from internal and/or external data files) as required by that subroutine. Cross-talk between subroutines is limited t o interrogation, and any answer which causes a change in an earlier-computed parameter results i n a recalculation of t h a t parameter. 

The paper includes logic to aid in sizing of various engine requirements.},

keywords = {14. Weight Engineering - Missile Design},

pubstate = {published},

tppubtype = {inproceedings}

}

@inproceedings{1048,

title = {1048. Flight Test of a Spin Parachute for Use With a Super Arcas Sounding Rocket},

author = {M N Silbert},

url = {https://www.sawe.org/product/paper-1048},

year  = {1975},

date = {1975-05-01},

booktitle = {34th Annual Conference, Seattle, Washington, May 5-7},

pages = {22},

publisher = {Society of Allied Weight Engineers, Inc.},

address = {Seattle, Washington},

abstract = {A specially configured 16.6-foot (5.1-meter) Disc Gap Band (DGB) Spin Parachute has been designed, 

developed, integrated with a modified Super Arcas launch vehicle, and qualified by flight testing. 

The total payload weight, which includes the Spin Parachute and a scientific payload, was 17.6 

pounds (7.99 kg). Total liftoff weight was 100.3 pounds (45.54kg). The Super Arcas vehicle was 

despun from 18.4cps, 127.3 seconds after launch. Payload separation occurred one second later at 

an altitude of 244,170 feet (74.4km). The Spin Parachute and payload attained a maximum spin rate 

of 2.4cps approximately 97 seconds after payload separation. The total suspended weight of the Spin 

Parachute and payload was 14.64 pounds(6.65kg). 

This Super Arcas vehicle contained several innovations which were necessary to provide mission support 

for the scientific payload. These included the first despin of the Super Arcas vehicle and the first 

high-altitude deployment of a Spin Parachute.},

keywords = {14. Weight Engineering - Missile Design},

pubstate = {published},

tppubtype = {inproceedings}

}

@inproceedings{0928,

title = {928. Aspects of Dynamic Balance of High RPM Spin Stabilized Rockets},

author = {R Schliesser},

url = {https://www.sawe.org/product/paper-0928},

year  = {1972},

date = {1972-05-01},

booktitle = {31st Annual Conference, Atlanta, Georgia, May 22-25},

pages = {20},

publisher = {Society of Allied Weight Engineers, Inc.},

address = {Atlanta, Georgia},

abstract = {The first part of the presentation introduces the Emerson ANSSR rocket, its mission, physical description, and firing sequence. It also shows the relative importance of proper dynamic balance. 

The second part deals with the types of balance mechanisms employed and discuss their general merits. Also discussed are the balancing machines used during the recent ANSSR program and the value of component balancing. 

The third part of the paper discusses peculiarities of certain balancing machines, in particular, fixed rate machines. It questions the use of fixed rate machines when used with an experimental, complex workpiece.},

keywords = {14. Weight Engineering - Missile Design},

pubstate = {published},

tppubtype = {inproceedings}

}

@inproceedings{0741,

title = {741. Influences of Dynamic Unbalance of Spin Stabilized Rockets},

author = {R I Figueras},

url = {https://www.sawe.org/product/paper-0741},

year  = {1969},

date = {1969-05-01},

booktitle = {28th Annual Conference, San Francisco, California, May 5-8},

pages = {22},

publisher = {Society of Allied Weight Engineers, Inc.},

address = {San Francisco, California},

abstract = {The accuracy of a spin stabilized rockets very sensitive to dynamic unbalance. This paper presents a description of the mechanics of missile dispersion due to dynamic unbalance and, also, discusses some of the problems associated with achieving a balanced spin stabilized rocket.},

keywords = {14. Weight Engineering - Missile Design},

pubstate = {published},

tppubtype = {inproceedings}

}

@inproceedings{0593,

title = {593. The Potential of a State-Of-The-Art Recoverable Launch Vehicle},

author = {J M Youngs},

url = {https://www.sawe.org/product/paper-0593},

year  = {1967},

date = {1967-05-01},

booktitle = {26th Annual Conference, Boston, Massachusetts, May 1-4},

pages = {15},

publisher = {Society of Allied Weight Engineers, Inc.},

address = {Boston, Massachusetts},

abstract = {In recent years there has been increased interest in the possible use of reusable launch vehicles. The economics of using a launch vehicle over and over certainly makes it a very attractive idea. At present, literally tons of exceedingly expensive hardware is abandoned in space or allowed to burn up upon re-entry into the earth's atmosphere. One possible way of avoiding this extremely wasteful procedure is to provide a launch vehicle with a crew and flyback capability. 

This paper describes a launch vehicle with this capability that has been developed at a conceptual level using state-of-the-art designs and materials. It is a vertical takeoff concept with flyback and horizontal landing capability. Basically, it represents the fusion of both rocket and airplane design considerations. This unusual configuration has some very interesting and unique problems associated with its design and mass properties control.},

keywords = {14. Weight Engineering - Missile Design},

pubstate = {published},

tppubtype = {inproceedings}

}

@inproceedings{0581,

title = {581. Preliminary Design Method for Weight Optimization},

author = {E C Stevens},

url = {https://www.sawe.org/product/paper-0581},

year  = {1966},

date = {1966-05-01},

booktitle = {25th Annual Conference, San Diego, California, May 2-5},

pages = {32},

publisher = {Society of Allied Weight Engineers, Inc.},

address = {San Diego, California},

abstract = {Rapidly changing design technology and development of new structural materials necessitate development of high speed computer programs to account for these advances in the preliminary design phase. A standard payoff function used in many current aerospace designs is weight minimization. Although this criteria does not necessarily yield an 'optimum' design (including cost effectiveness), it is quite valuable in determining performance effectiveness and often is at least an indicator of optimum design trends. 

The technique described in this paper has been successfully applied as a preliminary design method for minimizing the gross weight of a solid rocket propulsion system subject to performance, geometric, and design constraints which are obtained from operational or environmental restrictions. This technique utilizes the method of undetermined multipliers with a modified Newton-Raphson iteration process. 

The sample problem was formulated by developing an overall trade-off study mathematical model consisting of several environmental sub- models and a gross weight minimization model. The environmental submodels are used to define the in-flight loading and heating environment and provide design constraints for the gross weight minimization model. 

The subject method has been shown to be adequate for use as: 

- A preliminary design tool for establishing optimum design parameters 

- A design tool for comparative evaluation of structural materials and design concepts 

- A method for definition of areas where component research and development should be most fruitful 

This method has also been used to demonstrate the concept that to obtain maximum benefit from weight minimization, each structural material must be allowed to operate with its own set of optimum design parameters. 

In addition to a general description of the weight optimization method, a sample problem of multistage solid propellant rocket is shown, including details of methodology, types of problems encountered in the optimization process, and sample results. This problem is formulated as a gross weight minimization problem with burn-out velocity, case design, overall length, and burn-out acceleration constraints. Application of this method to turbine engine optimization in aircraft mission analysis is also discussed.},

keywords = {14. Weight Engineering - Missile Design},

pubstate = {published},

tppubtype = {inproceedings}

}

@inproceedings{0585,

title = {585. The Graphical Solutions of Oblate Spheroid Shell Intersection},

author = {J M Youngs},

url = {https://www.sawe.org/product/paper-0585},

year  = {1966},

date = {1966-05-01},

booktitle = {25th Annual Conference, San Diego, California, May 2-5},

pages = {22},

publisher = {Society of Allied Weight Engineers, Inc.},

address = {San Diego, California},

abstract = {When elliptical bulkheads are used for cylindrical tank ends, certain problems arise in relation to the calculation of exposed surface area and enclosed volume of the tank ends. The case of a single bulkhead resolves itself into the solution of the surface area and volume of an oblate spheroid, for the bulkhead is essentially an oblate spheroid shell that has been split in half by the plane of the major axes. The solution of the problem is readily available and can be expressed in fairly simple mathematical terms. 

A more serious and complicated problem arises when two tanks are clustered in such a way that the walls and bulkheads intersect. This condition results whenever the distance between the longitudinal axes of the tanks is less than the diameter of the individual tanks. Tanks would be arranged in this fashion to obtain better volume utilization within a given structural envelope. 

This paper deals with a graphical approach to determining surface areas and volumes of intersecting elliptical bulkheads. A rigorous mathematical solution of the problem requires very complicated calculus techniques and the solution resolves itself into elliptic integral power series. Even assuming that one is familiar with the techniques Involved, the solution for even only one set of conditions is exceedingly time consuming and because of the difficulty the chance of error high. The use of the curves derived in this paper reduces the problem to one of extreme simplicity. Someone without the slightest awareness of the complexities of the problem can obtain exact answers for the volume and answers within ~ 3 % for the surface area of intersecting elliptical bulkheads. The total time required to do this is mere minutes.},

keywords = {14. Weight Engineering - Missile Design},

pubstate = {published},

tppubtype = {inproceedings}

}

@inproceedings{0478,

title = {478. Monte Carlo Techniques as Applied to the A3 Polaris Missile},

author = {R B Heffron and L G Lauger and R W Harrell},

url = {https://www.sawe.org/product/paper-0478},

year  = {1965},

date = {1965-05-01},

booktitle = {24th Annual Conference, Denver, Colorado, May 17-19},

pages = {17},

publisher = {Society of Allied Weight Engineers, Inc.},

address = {Denver, Colorado},

abstract = {A concept called 'performance reliability' was used in designing the Polaris A3 missile. This concept allows a low percentage of subsystem failures in exchange for increased missile system performance (range or payload). This approach is necessary since the size of a submarine launch tube places restrictions on missile sizing. The only way to increase payload or range without changing launch tube size, therefore, is to improve state of the art, or accept reduced reliability. Naturally, both methods were used in designing the A3, however this paper will discuss only the method of reducing reliability to reduce weight in the second stage flight control subsystem. Once it has been decided to reduce reliability, a method must be used to actually determine analytically the probability of success of a subsystem. 

The analytical method finally decided upon was an iterative computer technique using random numbers. This technique is commonly called 'Monte Carlo.' The paper will discuss, within the limitations of security, the weight saving of a Monte Carlo designed fluid injection thrust vector control system having a small probability of failure vs. a system designed for worst on worst conditions. 

The success of the Monte Carlo Technique in the thrust vector control system led to serious consideration of its use in other areas which are primary weight responsibilities. For example, the tolerances on the c.g.'s and M.I.'s are very tedious to find in closed form, but very easy to find using Monte Carlo - if computer time and talent are available. Another example is the tolerance on weights during burning. The weighed inert weight of a motor includes a certain amount of burnable inert weight, as well as fixed inert weight. Since they are not independent variables, then the problem of finding tolerances on inert weight near burnout of a stage becomes a problem which responds nicely to Monte Carlo, whereas the root-sum-square technique is invalid because it assumes independence of variables. 

These cases are discussed, have been used in Polaris analysis, and have provided reasonable answers. Problems arising in their use are also discussed. It is felt that Monte Carlo will come into increased use as an analytical tool.},

keywords = {14. Weight Engineering - Missile Design},

pubstate = {published},

tppubtype = {inproceedings}

}

@inproceedings{0272,

title = {272. Introduction to Solid Propellant Rockets},

author = {J O Crum},

url = {https://www.sawe.org/product/paper-0272},

year  = {1961},

date = {1961-05-01},

booktitle = {20th National Conference, Akron, Ohio, May 15-18},

pages = {16},

publisher = {Society of Allied Weight Engineers, Inc.},

address = {Akron, Ohio},

abstract = {This paper was presented at the Twentieth Annual National Conference of the Society of Aeronautical Weight Engineers at Akron, Ohio, May 15-18, 1961. The paper is intended to provide an introduction for those not familiar with the field of solid propellant rockets. It provides a basic overview of how a solid propellant rocket works, and discusses types of propellants, propellant grain configuration, thrust vector control, and thrust termination. The state-of-the-art for solid propellant rockets in 1961 is also discussed.},

keywords = {14. Weight Engineering - Missile Design},

pubstate = {published},

tppubtype = {inproceedings}

}

@inproceedings{0274,

title = {274. Aspects of High Performance Sounding Rockets},

author = {S L Tuttle},

url = {https://www.sawe.org/product/paper-0274},

year  = {1961},

date = {1961-05-01},

booktitle = {20th National Conference, Akron, Ohio, May 15-18},

pages = {9},

publisher = {Society of Allied Weight Engineers, Inc.},

address = {Akron, Ohio},

abstract = {This paper was presented at the Twentieth National Conference of the Society of Aeronautical Weight Engineers at Akron, Ohio, May 15 - 18, 1961. A philosophy for the design of high performance sounding rockets as evolved by personnel of the University of Michigan, Aircraft Propulsion Laboratory, is presented. The term sounding rocket is generally taken to mean a vehicle capable of a vertical probe in the altitude region of a few miles to a few thousand miles. High performance implies the capability of exceeding 1,000 miles. 

Stress is laid on the fundamental differences in performance criteria of this type of research vehicle, and one designed primarily for military use. As the sounding rocket is primarily a research vehicle, it should be inexpensive, reasonably simple, and relatively versatile. 

These requirements have led to the conclusion that the most effective and least expensive high performance sounding rocket will be a single stage, non-cryogenic liquid (ideally a mono-propellant) propelled system with a gross loaded weight of two - three tons. 

By way of illustration, a basic design and the considerations which led to this design are presented. It is concluded that it is within the present state of the art to build a sounding rocket capable of carrying a 25 pound payload above 1,500 miles, at a cost of less than $50,000 per vehicle. 

Finally, by extending the state of the art somewhat and by relaxing some of the previous requirements, the possibility of a single stage rocket, capable of putting small payloads above 4,000 miles, is demonstrated. The energy requirement for a 4,000 mile vertical probe is equivalent to that for a low satellite.},

keywords = {14. Weight Engineering - Missile Design},

pubstate = {published},

tppubtype = {inproceedings}

}

@inproceedings{0275,

title = {275. Some Reasons for Concern About the ''M'' in F=MA in Rocketry},

author = {G R Blayzor},

url = {https://www.sawe.org/product/paper-0275},

year  = {1961},

date = {1961-05-01},

booktitle = {20th National Conference, Akron, Ohio, May 15-18},

pages = {6},

publisher = {Society of Allied Weight Engineers, Inc.},

address = {Akron, Ohio},

abstract = {This paper was presented at the Twentieth National Conference of the Society of Aeronautical Weight Engineers at Akron, Ohio, May 15 - 18, 1961. This paper will attempt to analyze one of the most severe problems facing the weight specialist in the field of rocketry. The problem is the continuing and significant need for answering questions as to the importance of weight knowledge, or weight control. 

The prevalence of such questions is discussed, in an attempt to explore the reasons for any mystery about the weight parameter. The writer will try to pinpoint some of the types who most frequently raise questions of this nature, and will point out that the reasons are in many instances very legitimate. The weight specialist is frequently confronted by demands for very specific answers where such answers do not exist. All too often the lack of ability to realistically satisfy such demands, or to present a clear cut case procreates an even greater conviction that missile weight is indeed, a minor problem and of no particular consequence to the given program. 

The writer will discuss some of the more fundamental reasons which are not always obvious, but which point up the need for adequate knowledge of weight parameters. 

The areas discussed will include cost of overdesign versus weight control conservatisms required in mission planning due to weight uncertainties and the fallacy of miniaturizing components while oversizing the vehicle. 

The paper will conclude with a broad attempt to point out the obligations of the missile weight specialist with respect to this problem area. The fact that the weight engineer is guilty of poor salesmanship and the obligations with which he is confronted are summarized. The writer will attempt to get himself shot by indicating some of those areas in which a weight specialist must excel if he ever wishes to see his profession placed in the technical realm of respect which it deserves.},

keywords = {14. Weight Engineering - Missile Design},

pubstate = {published},

tppubtype = {inproceedings}

}

@inproceedings{0276,

title = {276. Weight in Overall Missile Performance},

author = {J T Blake},

url = {https://www.sawe.org/product/paper-0276},

year  = {1961},

date = {1961-05-01},

booktitle = {20th National Conference, Akron, Ohio, May 15-18},

pages = {11},

publisher = {Society of Allied Weight Engineers, Inc.},

address = {Akron, Ohio},

abstract = {This paper was presented at the Twentieth National Conference of the Society of Aeronautical Weight Engineers at Akron, Ohio, May 15 - 18, 1961. This paper describes the problems of rocket and space weights and weight control from the standpoint of the user of this information. 

One of the major tasks in the Atlas Program is the precision determination of the over-all Weapon System performance: this is the determination of the range to which one is willing to assure SAC that they can safely target the missile. This range is referred to in the Atlas Program as the usable range. It should be emphasized that the usable range is not the average or maximum distance that the rocket will go, but the distance that you dare attempt to make it go.},

keywords = {14. Weight Engineering - Missile Design},

pubstate = {published},

tppubtype = {inproceedings}

}

@inproceedings{0216,

title = {216. Design of Take-Off Masses of Multiple Stage, ICBM-Type Rockets With Given Range and Payload Objectives},

author = {H W Baer},

url = {https://www.sawe.org/product/paper-0216},

year  = {1959},

date = {1959-05-01},

booktitle = {18th National Conference, Henry Grady Hotel, Atlanta, Georgia, May 18-21},

pages = {25},

publisher = {Society of Allied Weight Engineers, Inc.},

address = {Atlanta, Georgia},

abstract = {This paper discusses design parameters which determine the take-off masses of multiple-stage rockets and the gross and burnout masses of their individual stages. An effective mass ratio of the rocket is defined, which takes into account the finite burnout masses of the stages as well as the effect of the payload mass on each stage. This effective mass ratio is shown to be optimally expressible in terms of the number of stages, a payload ratio and a structural factor, provided that the structural factor is not a function of the stages. For given range and payload mass of a rocket, the chosen parameters enable the prediction of necessary take-off mass and required number of stages. Several example problems are worked out, which illustrate a procedure for obtaining take-off mass and individual stage masses of ICBM-type rockets. The quantities which determine these masses are: initial thrust-to-weight ratio, specific impulse of the propellants, payload mass, range and realizable structural factor. The dependence of the take-off mass as well as the achievable range on the realizable structural factor is discussed quantitatively.},

keywords = {14. Weight Engineering - Missile Design},

pubstate = {published},

tppubtype = {inproceedings}

}