SAWE Technical Papers

@inproceedings{3710,

title = {3710. Application of the Law of Propagation of Uncertainties to a Weight and CG},

author = {Anjie Emmett},

url = {https://www.sawe.org/product/paper-3710},

year  = {2019},

date = {2019-05-01},

booktitle = {78th Annual Conference, Norfolk, VA},

pages = {49},

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

address = {Norfolk, Virginia},

abstract = {In order to quantify potential errors in a measurement system, the uncertainties of all measurement sources must be combined to generate a total system uncertainty. This quantified measurement system uncertainty may be used as a decision-making tool to determine the required accuracy of measurement devices such as load cells, scales, and laser trackers.For NASA's Ascent Abort 2 (AA-2) Flight Test, such an uncertainty quantification was performed to ensure that the Ground Support Equipment (GSE) designed to measure the mass and center of gravity (CG) of a Crew Module (CM) would meet the accuracy requirements set forth by the program. The uncertainties of the load cells used were combined with the laser tracker system's positional uncertainty to determine the overall measurement system uncertainty, which met program requirements.},

keywords = {15. Weight Engineering - Missile Estimation, 21. Weight Engineering - Statistical Studies},

pubstate = {published},

tppubtype = {inproceedings}

}

@inproceedings{3701,

title = {3701. Mass Properties in Support of Class Analysis (a.k.a. MP End of Days)},

author = {Ricardo Roy},

url = {https://www.sawe.org/product/paper-3701},

year  = {2018},

date = {2018-05-01},

booktitle = {77th Annual Conference, Irving, Texas},

pages = {18},

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

address = {Irving, Texas},

abstract = {You receive the following Program Office Request: In lieu of mission specific analysis, it is dictated to you that Class Analysis be instituted to lower overall program cost. There may be other methods to address this request but this paper addresses one process that provided mass properties in support of program 'Class Analysis'. To get started, a definition of Class Analysis is prudent. Class Analysis is any single study that incorporates all conceived configurations of a vehicle from mass properties (MP) perspective and the MP uncertainties associated with them.The main purpose is to provide a range of mass properties with a high likelihood that the current and future fleet elements will not exceed them.This process is based on simulated aerospace hardware data and does not reflect any specific line of vehicles. Additionally, the same process may be applied to non-aerospace production programs. All that is required is a good history of launch vehicle segments along with mass properties and uncertainties for each segment. Ten (10) years of history is ideal, but a smaller term is acceptable knowing that risk may be incurred. A segment is defined as any portion of the launch vehicle that may get jettisoned during the launch cycle. A robust verification process to validate that the assumed variables are still compliant is also a must.This analysis will not only be performed by the mass properties group but also by all the downstream users (Guidance, Flight Mechanics, Separation, Structures, Ground Ops ...etc.). Mass Properties is the initial cog in a long string of analysis that will be scrutinized. This being said, the mass properties group must not operate in a vacuum, but coordinate with these downstream users to access the effects of your assumptions. All data units herein will be presented as: Mass (M) = pounds-mass (lbm); Center of Mass (CM) = inches (in) or stations; and Moment and Product of Inertias (MOI and POI) = slug-foot2 (sl-ft2). Inertias will be in respect to the CM. A positive integral definition will be used for POI.},

keywords = {15. Weight Engineering - Missile Estimation, 19. Weight Engineering - Spacecraft Estimation, 21. Weight Engineering - Statistical Studies},

pubstate = {published},

tppubtype = {inproceedings}

}

@inproceedings{2176,

title = {2176. Sizing Solid Propellant Boosters for Weight Growth},

author = {G T Richbourg and L M Stuntz},

url = {https://www.sawe.org/product/paper-2176},

year  = {1993},

date = {1993-05-01},

booktitle = {52nd Annual Conference, Biloxi, Mississippi, May 24-26},

pages = {14},

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

address = {Biloxi, Mississippi},

abstract = {A major contributing factor in sizing a solid propellant booster is the payload or throw weight, Usually, this weight includes an allowance for growth to allow for historical weight growth trends, Yet, another factor sometimes overlooked, is a weight growth allowance for the booster itself, Historical growth trends show booster weight also changes with program maturity. This usually means increasing booster inert weight and decreasing booster propellant weight. This paper gives one approach to sizing a booster that will meet all program mission performance requirements with a design throw weight even when the booster inert weight increases and propellant weight decreases.},

keywords = {15. Weight Engineering - Missile Estimation},

pubstate = {published},

tppubtype = {inproceedings}

}

@inproceedings{1992,

title = {1992. Launch Vehicle Synthesis and Methodology},

author = {R S Pribnow},

url = {https://www.sawe.org/product/paper-1992},

year  = {1991},

date = {1991-05-01},

booktitle = {50th Annual Conference, San Diego, California, May 20-22},

pages = {12},

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

address = {San Diego, California},

abstract = {Traditional approximate methods for sizing have been replaced at General Dynamics Space Systems (GDSS) with more exact weight and geometry modeling of vehicles. Vehicle models consisting of series of algebraic equations describing the vehicle geometry and weight have been utilized in two ways. - A program called SIZIT utilizes the weight models and the rocket equation method to size expendable launch vehicles with a high degree of versatility and accuracy. - A program called FASTPASS uses the weight models to optimally size a vehicle while running a trajectory simulation. This eliminates several steps in the vehicle synthesis process and results in a vehicle that meets performance requirements at minimum vehicle cost.},

keywords = {15. Weight Engineering - Missile Estimation},

pubstate = {published},

tppubtype = {inproceedings}

}

@inproceedings{1900,

title = {1900. Launch Vehicle Synthesis and a Computer Modeling Approach},

author = {A Prast and J Szedule},

url = {https://www.sawe.org/product/paper-1900},

year  = {1989},

date = {1989-05-01},

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

pages = {13},

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

address = {Alexandria, Virginia},

abstract = {Launch vehicle synthesis has evolved over the past four years at General Dynamics Space Systems Division (GDSS) from use of the basic rocket equation to detailed weights, propulsion, and trajectory computer modeling. This paper discusses the development of various methods of vehicle synthesis at GDSS and how this work has led to the computer synthesis program now called Flexible Analysis for Synthesis, Trajectory, and Performance on Advanced Space Systems (FASTPASS).},

keywords = {15. Weight Engineering - Missile Estimation},

pubstate = {published},

tppubtype = {inproceedings}

}

@inproceedings{1525,

title = {1525. Preliminary Design Analysis of Missile Weight, Cost and Schedule Interactions and Risk Assignments},

author = {M Skrtic and R S St. John},

url = {https://www.sawe.org/product/paper-1525},

year  = {1983},

date = {1983-05-01},

booktitle = {42nd Annual Conference, Anaheim, California, May 23-25},

pages = {22},

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

address = {Anaheim, California},

abstract = {In the preliminary design of a hardware item, it is often necessary to evaluate alternative design approaches, 

materials and capabilities that permit future product improvements while maintaining basic performance 

characteristics. Due to the complex nature of modern equipment and the extreme environments specified, these 

choices impact end-item cost, delivery schedule and,in some cases, performance, in a positive or negative 

manner. Frequently, these choices are made considering only one or two criteria, i.e., a material change for 

a particular component to reduce overall weight or to increase component lifetime. Only later is it discovered 

that this change requires procurement of scarce material, new tooling and cutting equipment and, perhaps, 

development of new manufacturing process specifications. In the final analysis, the end item will weigh less 

and have a greater service lifetime, but its delivery schedule may be incompatible with customer requirements 

and its cost growth may preclude quantity purchases. As the complexity of the item grows, the more decisions 

and interactions there are and the more rapidly the process of evaluating product-improvement options outstrips 

an individual's capability to assess the final impact on cost, schedule, performance and program risk. 

In developing a missile system that incorporates many interacting subsystems, timely evaluation of all options 

and selection of the path that provides the least cost and schedule impact while meeting requirements is, at 

best, difficult. As the use of computers has become widespread, many sophisticated techniques have been 

developed to evaluate alternatives; however, the amount of input data required to use these tools has also grown. 

As a result, few (if any) available tools lend themselves to use in preliminary design or proposal efforts where 

the demand is for 'instant results.' 

This paper discusses a methodology for evaluating alternative strategies to incorporate requirement change or 

product-improvement alternatives during the proposal or preliminary design stage of missile development. The 

entire analysis was performed on an Apple II Plus* computer. In addition to several specialized personal routines, 

the following commercially available software was used: 

o Hierographic Transport - Product of GSR Associates 

o Apple Plot - Product of Apple Compute rInc. 

o VisiCalc - Product of VisiCorp Personal Software 

o VisiPlot - Product of VisiCorp Personal Software 

o AMPERGRAPH - Product of MADWEST Software 

o TASC compiler - Product of Microsoft Corporation 

o Apple Writer - Product of Apple Computer Inc.},

keywords = {15. Weight Engineering - Missile Estimation},

pubstate = {published},

tppubtype = {inproceedings}

}

@inproceedings{1497,

title = {1497. Missile Body Weight Prediction},

author = {J R Atkinson and R N Staton},

url = {https://www.sawe.org/product/paper-1497},

year  = {1982},

date = {1982-05-01},

booktitle = {41st Annual Conference, San Jose, California, May 17-19},

pages = {63},

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

address = {San Jose, California},

abstract = {Vought has developed a technique for predicting missile body weight which combines an optimum shell analyzed for applied loads and material usage with weight penalties added for specific design features. The optimum shell is sized for buckling within a given section. Cylindrical and conical body sections are analyzed. Allowable buckling stress methods are taken from NASA CR912 'Shell Analysis Manual.' I Applied stress level s are computed using standard analysis expressions. Stress ratios are computed and the section is sized for the 1imiting value of the 1oading conditions considered. This rough sizing is further refined by consideration of combined loading effects. Penalty weights computed using a combination of analytical and empirical techniques include: - Weight penalty for minimum gauge - Weight penalty for handling - Weight penalty for joints - Pressure bulkhead weight - Wing and tail attachment - Weight penalty for access doors - Weight penalty for air launch features - Boat tail close out ring - Fastener and miscellaneous a1lowance. The method presented in this paper has been expanded, adapted for computer execution and incorporated into Vought's Missi1e Integrated Design and Analysis System, MIDAS.},

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

keywords = {15. Weight Engineering - Missile Estimation, Mike Hackney Best Paper Award},

pubstate = {published},

tppubtype = {inproceedings}

}

@inproceedings{1431,

title = {1431. Development of the Missile Integrated Design Analysis System (MIDAS), a Computurized Missile Synthesis for the 1980's},

author = {A J Roch and R N Staton},

url = {https://www.sawe.org/product/paper-1431},

year  = {1981},

date = {1981-05-01},

booktitle = {40th Annual Conference, Dayton, Ohio, May 4-7},

pages = {15},

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

address = {Dayton, Ohio},

abstract = {The Missile Integrated Design Analysis System (MIDAS) is being developed to fulfill the need 

for computerized missile synthesis at Vought. MIDAS brings together in one computer system the 

advanced design engineering analysis techniques and work procedures that produce sized and 

optimized missile systems. MIDAS can be used to analyze specific missile configurations or to 

quickly generate trade-off data in the form of missile weight,lethality, effectiveness, cost, 

etc., as functions of various performance and operational parameters. This paper summarizes 

the first fifteen months of the MIDAS development 

effort.},

keywords = {15. Weight Engineering - Missile Estimation},

pubstate = {published},

tppubtype = {inproceedings}

}

@inproceedings{0775,

title = {775. Predictive Design for Spacecraft Propulsion},

author = {W G Andrews and R P Gallant and D R Reed},

url = {https://www.sawe.org/product/paper-0775},

year  = {1969},

date = {1969-05-01},

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

pages = {21},

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

address = {San Francisco, California},

abstract = {In this paper predictive methods applicable to the design definition of a class of solid propellant rocket motors are discussed. These motors deliver from 10,000 to 750,000 Ibf-sec total impulse and have spherical case diameters from 12 to 40 inches. A computer program which provides a means of rapidly sizing spherical and extended spherical rocket motors is described. The program was developed by correlating the performance and weight statistics of qualified motors to obtain empirical equations relating historical data to significant motor design parameters. 

Other aspects of solid propellant rocket motor design, such as reproducibility and reliability, are treated in considerable detail. This is accomplished by citing demonstrated and achievable values of the more pertinent ballistic and mass properties design parameters.},

keywords = {15. Weight Engineering - Missile Estimation},

pubstate = {published},

tppubtype = {inproceedings}

}

@inproceedings{0777,

title = {777. Comparison of Large Solid Rocket Motors Optimized for Recurring Cost or Gross Weight},

author = {R Krueger},

url = {https://www.sawe.org/product/paper-0777},

year  = {1969},

date = {1969-05-01},

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

pages = {40},

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

address = {San Francisco, California},

keywords = {15. Weight Engineering - Missile Estimation},

pubstate = {published},

tppubtype = {inproceedings}

}

@inproceedings{0662,

title = {662. A Semi-Empirical Method for Propellant Tank Weight Estimation},

author = {L Willoughby},

url = {https://www.sawe.org/product/paper-0662},

year  = {1968},

date = {1968-05-01},

booktitle = {27th Annual Conference, New Orleans, Louisiana, May 13-16},

pages = {31},

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

address = {New Orleans, Louisiana},

abstract = {This report presents information on a semi-empirical method for obtaining a quick and realistic tank weight for nine common configurations. An empirical weight equation is included for both pressure-fed and pump-fed systems. The equation for the pressure-fed system is based on limited data, but does have the correct relationship to the turbopump-fed system equation. 

General formulae, based on tank geometry and pressure effects, by which a amore detailed tank weight analysis may be performed, are also presented. The information contained in a table of general formulae permits weights to be calculated for any tank configuration except torodial. Data necessary for determining tank design pressured based on propellant characteristics and flow rate, engine performance parameters, plumbing geometry, and missile trajectory is also included. The empirical weight increment is divided into a number of nonoptimum factors to permit a separate analysis of some of these items if sufficient data are available. Separate analysis of tank cylindrical sections may be performed by utilizing the included curves for pressure-stabilized cylindrical shells subjected to axial loads and bending moments.},

keywords = {15. Weight Engineering - Missile Estimation},

pubstate = {published},

tppubtype = {inproceedings}

}

@inproceedings{0693,

title = {693. Titan III Propellant Management Technique},

author = {R A Thomas and R W VandeKoppel},

url = {https://www.sawe.org/product/paper-0693},

year  = {1968},

date = {1968-05-01},

booktitle = {27th Annual Conference, New Orleans, Louisiana, May 13-16},

pages = {28},

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

address = {New Orleans, Louisiana},

abstract = {The quantity of usable propellant remaining when the entire usable amount of the other propellant has been expended is defined as outage. Since this outage represents a launch vehicle perfomance penalty, outage mininimization through propellant management is an essential part of propulsion system design. Propellant management in the Titan 111 family of launch vehicles is achieved by preflight analyses and subsystem calibrations rather than through the use of an inflight propellant utilization system. This method simplifies the flight system and thus minimizes potential problems with flight hardware. 

The analytical technique of propellant management primarily consists of matching the propellant 1oad to the average inflight mixtur eratio. The accuracy to which this match can be achieved is influenced by factors which affect loaded mixture ratio or burned mixture ratio, such as tank gas pressure-time history, propellant temperature, vehicle acceleration, and loading accuracy. Consequently, the analytical technique must account for thev ariations in the loaded mixture ratio and the burned mixture ratio.T his is done on a statistical basis by determining the possible deviations from nominal and combining the effects. 

The treatment of the many variables that affect load and outage is discussed, together with flight results. It is shown that the Titan 111 propellant management technique has been effective in utilizing the usable propellant and minimizing outage.},

keywords = {15. Weight Engineering - Missile Estimation},

pubstate = {published},

tppubtype = {inproceedings}

}

@inproceedings{0599,

title = {599. Preliminary Weight Estimation of Liquid Propellant Stages},

author = {D R Pence},

url = {https://www.sawe.org/product/paper-0599},

year  = {1967},

date = {1967-05-01},

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

pages = {22},

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

address = {Boston, Massachusetts},

abstract = {The technique presented consists of a series of empirical equations which will allow stage weight, for any propellant combination, mixture ratio, and engine thrust level, to be estimated to a reasonable degree of accuracy using simple and readily available input parameters. 

The method has proven useful for 'first cut' estimates as well as a gross check on the validity of weights derived by more sophisticated analyses.},

keywords = {15. Weight Engineering - Missile Estimation},

pubstate = {published},

tppubtype = {inproceedings}

}

@inproceedings{0610,

title = {610. Structural Weight Estimation Methods for Small Air Launched Missiles},

author = {G R Williams},

url = {https://www.sawe.org/product/paper-0610},

year  = {1967},

date = {1967-05-01},

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

pages = {57},

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

address = {Boston, Massachusetts},

abstract = {Methods are developed for estimating the structural weight of air- to-air and air-to-ground missiles for preliminary design studies. The weight of each major structural component is derived as the s of 'basic weight' plus 'penalty weight.' Basic weight consists of the weight of structure required to satisfy ruggedness criteria and design loads. Penalty weight is the weight in excess of basic weight that is required for such functions as local load introduction, equipment support, and quick disconnects. The methods include consideration of component geometry, load, design temperature, material and type of construction. Structural weight of eight existing missiles are estimated by the developed methods. Comparison of these estimated weights to the actual weights indicates that the methods yield results of acceptable accuracy.},

keywords = {15. Weight Engineering - Missile Estimation},

pubstate = {published},

tppubtype = {inproceedings}

}

@inproceedings{0574,

title = {574. Ascent Aerodynamic Heating Effects in Advanced Design Weight Estimation},

author = {D L Schade},

url = {https://www.sawe.org/product/paper-0574},

year  = {1966},

date = {1966-05-01},

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

pages = {29},

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

address = {San Diego, California},

abstract = {This paper presents a computerized method of determining the external insulation requirements because of aerodynamic heating during the boost phase of flight on a rocket booster. 

The computer program's aim is to aid in weight estimating during the advanced design phase of engineering in a rapid and accurate fashion. Proficient use or the program does not require an extensive background in heat transfer. 

The basic requirement for input to this program has one outside dependent parameter. This parameter is that the vehicle fly a normal trajectory or that the trajectory data be based on a related booster vehicle. The trajectory data required are altitude and Mach number at the desired time interval. These data are generally related to a particular thrust-to-weight ratio. 

Calculation of the total heat transfer is done by numerical integration within the computer program. The resultant temperature and heat transfer values derived from the program are used to determine the required thickness of the insulation. 

The computed temperature is comparable to a mean booster R & D flight program temperature ~ 10%. 

The insulation discussed in this paper is based on a cork compound. Other types of insulation may be used by varying the proper input insulation characteristics. The computer program is presented so that it may provide a basis for insulation calculations and may be changed to fit any desired vehicle.},

keywords = {15. Weight Engineering - Missile Estimation},

pubstate = {published},

tppubtype = {inproceedings}

}

@inproceedings{0493,

title = {493. Weight Optimization of a Two Stage Reusable Orbital Carrier},

author = {R Jensen},

url = {https://www.sawe.org/product/paper-0493},

year  = {1965},

date = {1965-05-01},

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

pages = {32},

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

address = {Denver, Colorado},

abstract = {Man's desire to expand his knowledge of the solar system and the universe wiL. lead to a steadily increasing number of trips between the earth's surface and various orbits above our atmosphere. Technically, these trips are already fe:asible with either expendable or reusable boosters. However, at the increased launch rates expected within the next 10 to 20 years, the cost of expendable boosters becomes prohibitive. To remain within the cost constraint of that portion of the national budget allocated to space exploration, without s:?riously restricting future space travel, requires the development of efficient reusable launch systems. 

During the conceptual d::osign phase of these reusable syst2ms, weight and performance studies ar3 required to eliminate unpromising configurations and to assist in selecting and optimizing the most feasible designs. The objective of this paper is to present a means for performing these weight studies. 

A method is for rapidly determining the take-off weight and system dry weights of a two-stage reusable booster. The stage weights are described in quation form by deriving simplified expressions for the major functional components based on background experience and related technology. As baseline configurations are developed, these expressions are revised and expanded. As more detailed information is required, calculations may be greatly accelerated through the use of digital computers. Typfral equations for each stage are shown, including both all-rocket and rocket-airbreathing propulsion combinations for the first stage. Effects on stage and system weights caused by variations in sign parameters such as staging velocity and propulsion systems are presented. 

The two-stage booster which is used as an example has a LO2/RP-l fueled, winged first stage and a LO2/LH2 fueled, winged body second stage which also contains the payload. Tne ombined stages take off horizontally and, after completing its mission, each stage also lands horizontally.},

keywords = {15. Weight Engineering - Missile Estimation},

pubstate = {published},

tppubtype = {inproceedings}

}

@inproceedings{0425,

title = {425. A Method of Weight Estimation for Advanced Missile Design},

author = {G R Reitz},

url = {https://www.sawe.org/product/paper-0425},

year  = {1964},

date = {1964-05-01},

booktitle = {23rd National Conference / Sheraton, Dallas Hotel, Southland Center, Dallas, Texas May 18-21},

pages = {30},

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

address = {Dallas, Texas},

keywords = {15. Weight Engineering - Missile Estimation},

pubstate = {published},

tppubtype = {inproceedings}

}

@inproceedings{0383,

title = {383. Wire Weight Determination for Missiles},

author = {D L Goecks},

url = {https://www.sawe.org/product/paper-0383},

year  = {1963},

date = {1963-05-01},

booktitle = {22nd National Conference, St. Louis, Missouri, April 29 - May 2},

pages = {32},

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

address = {St. Louis, Missouri},

abstract = {This paper was presented at the Twenty-second Annual National Conference of the Society of Aeronautical Weight Engineers at St. Louis, Missouri, April 29-May 1, 1963. On any missile program one of the hardest items to try to determine at either an advanced design or hardware states is the weight of wire. The purpose of this report is to present a method that can be used to approximate wire weight for a missile that is yet in the engineer's mind and to determine accurate weights for wire on missiles that are in a hardware state. This causes two distinct and different situations to be covered. 

First, the Advanced Design problem. Missile wiring can be separated into three general categories: 

1) Instrumentation 

2) Guidance and Flight Control 

3) Power 

A base must be set. Before weight can be predicted for a theoretical missile, a weight must be determined for an existing missile. This existing weight is then turned into useful information by finding a common key to all missiles. 

What is a common key? First, look at Instrumentation. Overall, the amount of wire present depends solely on the number of measurements taken. Guidance and Flight Control depends solely on the number of measurements taken. Guidance and Flight Control depends upon the number of functions to be performed. Power depends on the loads necessary for flight. 

By breaking out Instrumentation pounds per measurement, Guidance and Flight Control pounds per function and Power pounds per amp-hr, the weight of an equivalent system can be determined. Factors then can be applied to take care of differences in length, diameter, configuration, etc. from the base missile.},

keywords = {15. Weight Engineering - Missile Estimation},

pubstate = {published},

tppubtype = {inproceedings}

}

@inproceedings{0385,

title = {385. Some Weight Aspects of a Missile Optimization Computer Program},

author = {W G Peterson},

url = {https://www.sawe.org/product/paper-0385},

year  = {1963},

date = {1963-05-01},

booktitle = {22nd National Conference, St. Louis, Missouri, April 29 - May 2},

pages = {22},

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

address = {St. Louis, Missouri},

abstract = {This paper was presented at the Twenty-second Annual National Conference of the Society of Aeronautical Weight Engineers at St. Louis, Missouri, April 29-May 1, 1963. The use of high-speed digital computers has become commonplace in the aerospace industry in recent years. One use is to facilitate the preliminary design of missiles through parametric analysis of various design characteristics. The sophistication of a particular program may be judged, from the weight engineer's point of view, by the approach taken to account for the missile's weight. This may vary from the use of a fixed mass fraction to the use of extremely detailed structural and weight analysis methods. 

This paper describes Program Omega, a solid propellant ballistic missile design and optimization method that takes advantage both of the capabilities of the computer and the judgment of the engineers in each of the many disciplines involved. Much of the program is devoted to the definition of propulsion and structural parameters to ensure the accuracy of the weight analysis. In order to illustrate the extent of the work undertaken in this approach, a rather brief outline of each program subroutine is presented along with a description of how these subroutines are integrated. Major emphasis is placed on the weight subroutine and other facets of the program where weight engineering is involved.},

keywords = {15. Weight Engineering - Missile Estimation},

pubstate = {published},

tppubtype = {inproceedings}

}

@inproceedings{0318,

title = {318. Weight Analysis of Ice and Frost Formation on the Liquid Oxygen Tank of a Missile},

author = {F Peters},

url = {https://www.sawe.org/product/paper-0318},

year  = {1962},

date = {1962-05-01},

booktitle = {21st National Conference, Seattle, Washington, May 14-17},

pages = {20},

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

address = {Seattle, Washington},

abstract = {This paper was presented at the Twenty-first Annual National Conference of the Society of Aeronautical Weight Engineers at Seattle, Washington, May 14-17, 1962. During the past few years, extensive testing has been performed with model tanks to determine the weight of ice and frost that might form on a liquid oxygen tank under varying climatic conditions. In extreme instances, weights as high as 1500 pounds in fair weather and 10,000 pounds in inclement weather were extrapolated from model test data. 

In a recent test, a camera was placed in an ICBM nose cone to photograph the booster during separation. These pictures clearly showed that 40 percent of the liquid oxygen tank was still covered with ice and frost at separation. It had previously been assumed that any ice and frost formed on the LO2 tank during filling would be shaken off by vibration, or melted away during boost through the atmosphere. 

Additional tests showed that model tank extrapolations were considerably higher than actual ice and frost formations observed on missiles. The test data were analyzed by standard statistical techniques, revealing the most important variables affecting ice and frost formation. These were found to be, in order of importance: hold period, ambient temperature, wind velocity, and relative humidity. The results obtained from this statistical analysis, when correlated with other test data, yield explanation for the ice and frost formation phenomena on a missile liquid oxygen tank.},

keywords = {15. Weight Engineering - Missile Estimation},

pubstate = {published},

tppubtype = {inproceedings}

}