affordable low fidelity target design competition final report

8/11/2019 Affordable Low Fidelity Target Design Competition Final Report

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2007/2008 AIAA MSTC ALFT Missile Design Competition

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Executive SummaryTarget vehicle systems are an essential element of the maturation and testing of the U.S.Ballistic Missile Defense (BMD) system. Physical tests utilizing such vehicles allow forthe evaluation of the BMD systems performance at a level that is beyond the capabilitiesof advanced computer modeling and more representative of an actual engagement.Drawbacks for testing with target missiles include the high costs incurred and thedifficulty of using targets to simulate a large variety of systems. Therefore, a new targetsystem that is both affordable and flexible is desired. The proposed Affordable LowFidelity Target (ALFT) system family is a low cost target that meets these needs.

A team of Georgia Institute of Technology students has conceptually designed a familyof ALFT systems in the AIAA/Missile Systems Technical Committee (MSTC) GraduateTeam Missile Design Competition during the 2007-2008 academic year that addressesthese perceived target drawbacks. In order to conceptually design this target family anappropriate process was developed. First, in-depth research into the problem, includingmotor and payload front section characteristics, was compiled from the open literature. AQuality Function Deployment (QFD) was also used to map the requirements to thesignificant engineering characteristics. This information allowed the team to generate alarge design space of vehicle alternatives. These options were mapped and amorphological analysis was conducted to identify the compatible options. The feasibledesigns were then evaluated using an extensive modeling and simulation environmentthat appropriately addressed the physics of all relevant disciplines for the target familys

performance. After this analysis, a family of vehicles was downselected from the feasibledesigns using multi-attribute decision making techniques. A higher fidelity design wasthen carried out on the most promising families of targets. Special attention was paid tothe design of the reentry object, including its thermal protection and propulsion systems.Safety, logistics, and support considerations were also addressed in the design.

The target family was designed with a preference of government furnished soundingrocket equipment to meet the needs of an assortment of missions because of their lowcost and asset availability. To further reduce the target family cost, the targets weredesigned to be unguided. This deviation from most targets currently in use eliminates theneed for a complex and costly active guidance system. This study focused on twodelivery orders (DOs): the first (DO1) requiring a minimum range of 1000 km, and thesecond (DO2) requiring a minimum range of 2500 km. Both DOs featured payloadsgreater than 400 kg. These payloads included a reentry object (RO) and an avionicssection (AS) which was designed to carry and deploy several associated objects (AOs).Two different ROs were designed: a non-separating conic (DO1) and a maneuverable bi-

conic with propulsive range extension capabilities (DO2).

Final downselection and design yielded a family solution to both DOs, with a commonfirst and final stage for each. The solution to DO1 consists of a Talos first stage and anM57 second stage, with a maximum range of 1488 km. The solution to DO2 consists of aTalos first stage, SR19 second stage, and M57 third stage, with a maximum range of4166.


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Table of ContentsExecutive Summary... ......................................................................................................... ii Table of Contents.... ........................................................................................................... iii List of Figures .. ................................................................................................................... v List of Tables .... ................................................................................................................ vii List of Acronyms ..... ........................................................................................................ viii Conceptual Design Team.. .................................................................................................. 1

Faculty Advisor.. ............................................................................................................. 1 ALFT Design Team Members and Responsibilities.. ..................................................... 1 Special Thanks.. .............................................................................................................. 1

Introduction.. ....................................................................................................................... 2 ALFT Overview.. ............................................................................................................ 2 The Need for Missile Defense .. ...................................................................................... 2 Request for Proposal.. ..................................................................................................... 5

Design Methodology.. ......................................................................................................... 7 Problem Definition.. ........................................................................................................ 7 Concept Selection .. ......................................................................................................... 9 Detailed Analysis.. .......................................................................................................... 9

Requirements Summary... ................................................................................................. 10 Identification of Concepts... .............................................................................................. 12 Modeling and Simulation Architecture... .......................................................................... 14

Propulsion ... .................................................................................................................. 15 Motor Database... ...................................................................................................... 15

Nozzle Diameter Determination ... ............................................................................ 15 Motor Characteristics Summary ... ............................................................................ 16

Geometry... .................................................................................................................... 16 Payload Details ... ...................................................................................................... 17 Launch Vehicle and Interstage Assumptions... ......................................................... 19

Aerodynamics ... ............................................................................................................ 19 Missile DATCOM ... ................................................................................................. 19 Missile Shape ... ......................................................................................................... 20

Trajectory... ................................................................................................................... 21 Boost Phase... ............................................................................................................ 22 Midcourse and Reentry Phases... .............................................................................. 23 Range Extension Phase ... .......................................................................................... 23 Code Evaluation... ..................................................................................................... 24

Thermal Analysis... ....................................................................................................... 25 Thermal Protection System Database ... .................................................................... 26

Zero-Order Conceptual Analysis... ........................................................................... 27 Higher Fidelity Analysis... ........................................................................................ 27

CAD... ........................................................................................................................... 28 Concept Selection ... .......................................................................................................... 30

Interactive Trade-off Tool... .......................................................................................... 30 Technique for Ordered Preference by Similarity to Ideal Solution... ....................... 30 Weighting Scenarios ... .............................................................................................. 31 Delivery Order Commonality ... ................................................................................ 32


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Architecture of the Tool... ......................................................................................... 32 Missile Downselection... ............................................................................................... 32

Missile Conceptual Design ... ............................................................................................ 39 Final System Design Overview ... ................................................................................. 39 Interstage Design ... ....................................................................................................... 39

Trajectory Performance ... ............................................................................................. 42 Thermal Protection System Design ... ........................................................................... 46 Front Section Design... .................................................................................................. 49

Launch Options... .............................................................................................................. 53 Shipping Logistics ... ..................................................................................................... 53 Air Launch ... ................................................................................................................. 54

Conclusions... .................................................................................................................... 55 Appendix A ... .................................................................................................................... 56 References... ...................................................................................................................... 68


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List of FiguresFigure 1 - Worldwide ballistic missile capabilities as of 1972.. ......................................... 3 Figure 2 - Worldwide ballistic missile capabilities as of 2005.. ......................................... 4 Figure 3 - BMD computational simulation and a physical target BMD test .. .................... 4 Figure 4 - Minuteman II missile system in flight .. ............................................................. 5 Figure 5 - Quality Function Deployment.. .......................................................................... 8 Figure 6 - IRMA tool ... ..................................................................................................... 13 Figure 7 - IRMA with compatibility constraints... ............................................................ 13 Figure 8 - Modeling and simulation architecture... ........................................................... 14 Figure 9 - Determining ATACMS exit area ... .................................................................. 15 Figure 10 - Geometry code operation ... ............................................................................ 17 Figure 11 - NASA bi-conic reentry object... ..................................................................... 17 Figure 12 - ALFT bi-conic reentry object... ...................................................................... 18 Figure 13 - ALFT conic reentry object... .......................................................................... 18 Figure 14 - Drag coefficient throughout the flight regime ... ............................................ 20

Figure 15 - Missile DATCOM shape definition... ............................................................ 20 Figure 16 - Trajectories for a notional missile... ............................................................... 21 Figure 17 Range Extension Trajectories for a notional missile... .................................. 24 Figure 18 - Code results vs. known data for Black Brant VC MK1... .............................. 25 Figure 19 - TPS database created for sizing and analysis... .............................................. 27 Figure 20 - Thermal M&S flow... ..................................................................................... 28 Figure 21 - Visualization of design space... ...................................................................... 29 Figure 22 - Parametric trade tool dashboard... .................................................................. 30 Figure 23 - Pareto frontier... .............................................................................................. 31 Figure 24 - Weighting scenarios... .................................................................................... 31 Figure 25 - Concept selection flowchart... ........................................................................ 32

Figure 26 - Proof of functionality ... .................................................................................. 34 Figure 27 - Best cost options ... ......................................................................................... 34 Figure 28 - Talos solid rocket motor... .............................................................................. 35 Figure 29 - SR19 and M57 solid rocket motors... ............................................................. 35 Figure 30 - Talos/M57, DO1 concept... ............................................................................ 37 Figure 31 - Talos/SR19/M57, DO2 concept... .................................................................. 38 Figure 32 - Starbird, illustrating Talos as first stage... ...................................................... 38 Figure 33 - Talos/M57 (DO1)... ........................................................................................ 41 Figure 34 - Talos/SR19/M57 (DO2)... .............................................................................. 41 Figure 35 - AS / M57 interstage drawing ... ...................................................................... 40 Figure 36 - DO1 trajectory - 1000km targeted (high trajectory) ... ................................... 44

Figure 37 - DO2 trajectory - maximum range ... ............................................................... 44 Figure 38 - DO1 mass change... ........................................................................................ 45 Figure 39 - DO2 mass change... ........................................................................................ 45 Figure 40 - Maximum surface temperatures during reentry for 2500 km ... ..................... 47 Figure 41 - TPS thickness required for reentry for 2500 km... ......................................... 47 Figure 42 - Maximum surface temperatures during for 4166 km... .................................. 48 Figure 43 - TPS thickness required for reentry for 4166 km... ......................................... 48 Figure 44 - LEROS- 1B ... ................................................................................................. 50


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Figure 45 - Bi-conic RO/AS breakout ... ........................................................................... 50 Figure 46 - Bi-conic RO/AS alternative view partially exploded... .................................. 51 Figure 47 - Conic RO/AS ... .............................................................................................. 51 Figure 48 - Reentry object and avionics section breakdown ... ......................................... 52 Figure 49 - Gravity air launch... ........................................................................................ 54

Figure 50 - Trapeze-lanyard air drop with parachute stabilization... ................................ 54


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List of TablesTable 1 - General delivery order specifications... ............................................................. 10 Table 2 - Reentry object specifications... .......................................................................... 10 Table 3 - Avionics section specifications ... ...................................................................... 10 Table 4 - Motors evaluated ... ............................................................................................ 11 Table 5 - Motor characteristics ... ...................................................................................... 16 Table 6 - Geometry mass estimation ... ............................................................................. 19 Table 7 - Cost data ... ......................................................................................................... 33 Table 8 - TOPSIS ideal solution preferences... ................................................................. 33 Table 9 - Launch angle sensitivity (DO1)... ...................................................................... 36 Table 10 - Transportability comparison (DO1) ... ............................................................. 36 Table 11 - Launch angle sensitivity (DO2)... .................................................................... 37 Table 12 - Transportability comparison (DO2) ... ............................................................. 37 Table 13 - RO Characteristics comparison ... .................................................................... 37 Table 14 - Talos/M57 (DO1) basic characteristics... ........................................................ 42

Table 15 - Talos/SR19/M57 (DO2) basic characteristics... .............................................. 42 Table 16 - Final geometry for interstages (DO1)... ........................................................... 39 Table 17 - Final geometry for interstages (DO2)... ........................................................... 39 Table 18 - Interstage mass estimates (DO1)... .................................................................. 40 Table 19 - Interstage mass estimates (DO2)... .................................................................. 40 Table 20 - Talos/M57 (DO1) trajectory performance... .................................................... 42 Table 21 - Talos/SR19/M57 (DO2) trajectory performance... .......................................... 43 Table 22 - Talos/SR19/M57 (DO2) TPS results... ............................................................ 46 Table 23 - Component requirements list from the TRD and SOW for DO1... ................. 49 Table 24 - Component requirements list from the TRD and SOW for DO2... ................. 49 Table 25 - Additional derived component requirements consist for DO2... ..................... 50

Table 26 - Front section component key... ........................................................................ 52


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List of AcronymsALFT Affordable Low Fidelity TargetASDL Aerospace Systems Design LaboratoryAO Associated ObjectAS Avionics SectionBMD Ballistic Missile DefenseCAD Computer Aided DesignCATIA Computer Aided Three Dimensional Interactive ApplicationCFP Contractor Furnished PropertyCONOPS Concept of OperationsCOTS Commercial Off-The-ShelfDO Delivery OrderDoD Department of DefenseEAFB Eglin AFBEMC Electromagnetic Compatibility

EMI Electromagnetic InterfaceES Experimental SubsystemsEWR Eastern and Western RangeFS Front SectionGFE Government Furnished EquipmentGFP Government Furnished PropertyHILM Hit Impact Location MeasurementIRMA Integrated Reconfigurable Matrix of AlternativesM&S Modeling and SimulationMADM Multi-Attribute Decision MakingMDA Missile Defense Agency

MDATCOM Missile Data CompendiumMMH Monomethyl HydrazineMON Mixed Oxides of NitrogenMSTC Missile Systems Technical CommitteeOTS Off-The-ShelfPAC-3 Patriot Advanced Capability-3PMRF Pacific Missile Range FacilityQFD Quality Function DeploymentRCS Reaction Control SystemRO Reentry ObjectRTS Reagan Test Site

SE Support EquipmentSOW Statement of WorkTMSS Thermal Management System SizerTOPSIS Technique for Order Preference by Similarity to Ideal SolutionTPS Thermal Protection SystemTRD Technical Requirements DocumentV&V Verification and ValidationVAFB Vandenberg Air Force Base


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VLDE Very Low Density ElastomericVSP Visual Sketch PadWFF NASA/Wallops Flight FacilityWMD Weapon of Mass DestructionWSMR White Sands Missile Range


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2007/2008 AIAA MSTC ALFT Graduate Missile Design Competition

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Conceptual Design TeamThe following proposal summarizes the work performed for the 2007-2008 AIAA MSTCMissile Graduate Design Competition. The conceptual design team for this competitionconsisted of graduate students and undergraduate students from the School of AerospaceEngineering at the Georgia Institute of Technology. Combined, these students contributedmore than 6,000 hours of analysis to the conceptual design during the period of technical

performance from September 1, 2007 through June 1, 2008.

Faculty Advisor Dr. Dimitri MavrisProfessor and Boeing Professor of Advanced Aerospace SystemsDirector, Aerospace Systems Design Laboratory (ASDL)Georgia Institute of Technology

ALFT Design Team Members and Responsibilities Mr. Adam Maser Program ManagerMr. Billy Gallagher Chief EngineerMr. Frank Coleman PropulsionMr. Lee Demory StructuresMr. Andrew Hensley PropulsionMr. Andrew Herron Thermal AnalysisMr. Kamal Kayat AerodynamicsMr. Brad Robertson Trajectory

Ms. Elizabeth Saltmarsh*

Preliminary DesignMr. Doug Stranghoener * VisualizationMr. Rob Willett * CAD

Special Thanks Mr. Robert Leginus Design Competition Subcommittee ChairMs. Rebecca Douglas Project AdvisorMr. Ian Stults Project AdvisorMr. Irian Ordaz Engineering Advisor

*Denotes Undergraduate Team Member


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IntroductionThe following proposal includes engineering analysis and hardware design associatedwith the buildup and launch of flight vehicles in support of ALFT missions. The designsolutions were created in accordance with the Statement of Work (SOW), TechnicalRequirements Document (TRD), December Kickoff Meeting discussions, and MarchSystems Requirements Review discussions. Concepts were explored with the knowledgethat the ALFT missions will be conducted from and staged out of test ranges including,

but not limited to, White Sands Missile Range (WSMR), the Reagan Test Site (RTS),Wake Island, Western Range, Pacific Missile Range Facility (PMRF), NASA/WallopsFlight Facility (WFF), Vandenberg Air Force Base (VAFB), and Eglin Air Force Base(EAFB). A focus has been given to ground launch, but air and sea launch capabilitieswere also explored. Also, because flight termination and active guidance sub-systems aredriving requirements for cost, alternative solutions were explored to eliminate the needfor these sub-systems.

ALFT Overview The ALFT system consists of a family of affordable non-separating and separating targetvehicles designed to complete two representative missions from 1000 to 2500 km inrange always using ground launch techniques, but also possessing air and sea launchcapabilities if possible. The payload of the ALFT system includes a maneuverable andrange extension capability RO and AS, with the ability to deploy AOs. The objective ofthe ALFT system is to provide a low cost, quick turn-around missile system that can beused for assessing and calibrating sensor system developments and modifications,

payload developments, sounding rocket experiments, and limited intercept experiments.

The Need for Missile DefenseThe current ballistic missile concern is different from the concern prevalent during theCold War. Wartime enmity causes state leadership to be more risk prone, and unstablegovernments can result in potential change in control of the military forces. Weapons ofmass destruction (WMD) are now a weapon of choice instead of a weapon of last resort,which is how they were viewed under the Cold War mindset. Antagonistic states want

ballistic missile technology to deter the United States or international intervention, and soWMDs are used to compensate for conventional strength. WMDs can also be used as ameans to coerce the United States and its allies. Missile defense serves as an enabler ofUnited States force projection.

The end of the Cold War has made [mutual assured destruction] largely irrelevant.Barely plausible when there was only one strategic opponent, the theory makes no sensein a multipolar world of proliferating nuclear powers. Mutual destruction is not likely towork against religious fanatics; desperate leaders may blackmail with nuclear weapons;

blackmail or accidents could run out of control. And when these dangers materialize, therefusal to have made timely provisions will shake confidence in all institutions of


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government. At a minimum, the rudiments of a defense system capable of rapidexpansion should be put into place.

- Henry Kissinger, March 9, 1995.

Figure 1 and Figure 2 show the worldwide ballistic missile capabilities as of 1972 and as

of 2005. These maps show in startling detail the degree of which ballistic missilecapability has spread throughout the world, from just a few nations to a major percentageof world nations. These figures illustrate further the need for a BMD system.

FIGURE 1 - WORLDWIDE BALLISTIC MISSILE CAPABILITIES AS OF 1972, WHERE NATIONS

IN ORANGE HAVE BALLISTIC MISSILE RANGES >2000 KM, AND NATIONS IN GREEN HAVEBALLISTIC MISSILE RANGES >1000KM [1]

The greatest strategic threat to the United States is an attack by one or more ballisticmissiles armed with nuclear or other weapons of mass destruction. Today, the UnitedStates remains vulnerable to this form of attack. Thus there is an urgent need for robustand layered missile defenses. Systems based on land, sea, air, and in space which arecapable of intercepting a missile during any phase of its flight are necessary to establish areliable defense.

As the United States Missile Defense Agency (MDA) advances with development anddeployment of its ballistic missile defense systems, a need is created to test and evaluatethese fast emerging systems. With interceptor missile capabilities proceeding at a rapid

pace, an ALFT system is needed for real threat simulation. There are two prominentforms of testing a missile defense system: simulations and computational analysis, andusing physical targets.


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FIGURE 2 - WORLDWIDE BALLISTIC MISSILE CAPABILITIES AS OF 2005, WHERE NATIONSIN ORANGE HAVE BALLISTIC MISSILE RANGES >2000 KM, AND NATIONS IN BLUE HAVE

BALLISTIC MISSILE RANGES >1000KM [1] [2]

The main benefits of using simulations and computational analysis is that they are cheap,repeatable, and efficient. Their major downfall is that they must by their nature makemany assumptions and use theoretical models, both of which introduce uncertainty.Physical targets are beneficial because they truly evaluate a systems performance inreality. However, they are hurt by a slow turnaround between tests and also high cost,which lower the number of available tests. Figure 3 below shows notional examples of acomputational simulation and a physical target test. Essentially, simulations are ideal forsizing and selection of preliminary designs, but a physical target will always be needed tovalidate a BMD system.

FIGURE 3 - LEFT: BMD COMPUTATIONAL SIMULATION USED BY THE ASDL; RIGHT: IMAGEOF A PHYSICAL TARGET BMD TEST [3]


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The most used existing target system is currently the Minuteman II motor. These are used because vendors have experience with the motors and there are large stockpiles of themotors. The system is however very expensive.

FIGURE 4 - MINUTEMAN II MISSILE SYSTEM IN FLIGHT [4]

Missile systems to be used as physical targets must be capable of providing targetmissiles with a range of 50 to 4000 km and be capable of flying various trajectories and

payloads. These systems have short lead times with relatively simple payloads. The mostcommon uses are as targets, experiment delivery vehicles and sensor systems test cuingobjects. The objective of the ALFT program is to provide low cost, quick turn-aroundmissile systems that can be used for assessing and calibrating sensor systemdevelopments and modifications, payload developments, sounding rocket experiments,and limited intercept experiments. Hence a new target system will play a critical role insafeguarding the United States and its allies in the twenty-first century.

Request for ProposalThe Affordable Low Fidelity Target Systems must be able to simulate a wide range of

potential threat vehicles over a large number of different ranges and boost regimes. Dueto the rapidly advancing progress and needs of the United States missile defenseoperations, the ALFT shall be a flexible, cost efficient and quick turn-around missilesystems. The capability of launch from non-terrestrial platforms, which may be executedvia either air or sea launch methods further defines the requirement for maximum


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flexibility. Thus, it shall also meet the specifications of existing ground facilities andgovernment furnished properties, such as WSMR and VAFB, for more traditional launchoperations. Regardless of launch mode, the ALFT shall be ready for launch within 20days of call-up up from the long-term storage condition. The system shall have acalculated launch target presentation availability of greater than 95% including the

reliability of the ALFT and support equipment in a variety of weather conditions.


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Design MethodologyAn appropriate process was developed to design the ALFT systems. This process wasdivided into three distinct phases, each of which concluded with an industry review.Phase 0 was the problem definition phase, which involved the background research intoBMD and targets of interest and the definition of the two requirements documents: theSOW and the TRD. Phase I was the concept selection phase. During this phase, modelingand simulation tools were created and used to select the family of ALFT target vehicles.Phase II was used to do the detailed analysis, including the design of the target missilefront section.

Problem DefinitionAt the beginning of the design process, the team began by conducting in-depth researchinto BMD systems and current target vehicles. This allowed the team to identify theshortfalls of current systems and the challenged faced in the design of the ALFT systems.After the background research was completed, the requirements documents, the SOWand TRD, were examined in detail and clarifications from the customer were obtained.The project plan and timeline as well as the modeling and simulation approach were alsodeveloped during this time.

Due to the high importance that the requirements place on the design of the reentry objectand the classified nature of reentry object data, a significant amount of research was thencarried out in order to further define the reentry object requirements. Also, therecommended government furnished motors and other motors of the same class wereresearched in an effort to determine their engineering characteristics.

At the Kickoff Meeting, which occurred at the conclusion to Phase 0, an interactive QFDwas presented, which allowed the team to gain valuable insight into the customerrequirements and target values. The QFD is a systems engineering tool that allows one tomap all of the customer requirements to the engineering characteristics. This helps toidentify the critical engineering characteristics in the design and the important tradestudies that must be conducted. The QFD developed for this project is shown in Figure 5 .As seen in the QFD, the cross range and down range distances were determined to be themost important engineering characteristics to consider, followed by launch vehicle lift-to-weight ratio and reentry object heat shielding.


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FIGURE 5 - QUALITY FUNCTION DEPLOYMENT


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Concept SelectionPhase I, the concept selection phase, began after the kickoff meeting. The goal of this

phase was to determine which combination of rocket motors was the best solution tothe ALFT design problem. First order modeling and simulation tools were developedduring this phase to address this. All of the relevant disciplines, including trajectory,geometry, aerodynamics, propulsion, and thermal, were addressed in this analysis. Next,these physics based tools were used to evaluate performance and identify feasibility ofeach concept for the two representative missions. The results of these modeling effortswere then used to select a family of ALFT missiles from the feasible options. This wasdone using an interactive Multi-Attribute Decision Making (MADM) tool with industryinput at a System Requirements Review that took place at the conclusion of Phase I.

Detailed AnalysisThe final phase of the design process, Phase II, focuses on the higher-fidelity analysis ofthe chosen concepts. Additionally, a key component of this phase was the design of thefront section of the missile. The composition and design of the reentry object andavionics section was laid out in detail. Logistics, support, and safety requirements, which

place constraints on such things as assembly, test, and launch, were also addressed in this phase. Details concerning the launch method, ship and shoot capability and missileconstruction were all examined. At the conclusion of the design, verification andvalidation will also be conducted on the ALFT concept to conform that it meets all of therequirements and specifications and performs as intended.


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Requirements Summary The problem is defined and the requirements are stated through the provided SOW andTRD documents, as well as from kickoff meeting and SRR discussions.

As noted in the overview section above, the primary goal of the ALFT system is to createlow-cost target systems. These systems will be capable of completing missions of varioustrajectories and payloads. These systems must allow for short lead times using an all upround concept of operations, where the complete missile stack can be delivered fullyassembled from the manufacturer. This ship and shoot method minimizes the assemblyand preparation time at the launch site.

Two specific DOs are defined for this project, a short-range mission (DO1) and a long-range mission (DO2). A comparison of the delivery orders is shown in Table 1 -Table 3 .

TABLE 1 - GENERAL DELIVERY ORDER SPECIFICATIONS

Parameter Delivery Order 1 Delivery Order 2Range 1000 km 2500+ km

Number of Stages 1-2 2-3Launch Options Ground and Sea Ground Only

TABLE 2 - REENTRY OBJECT SPECIFICATIONS

Parameter Delivery Order 1 Delivery Order 2

Type Non-Separating Separating

Shape Conic Bi-conic

Post ApogeeSurvival Altitude 100 km 40 km

Range Extension N/A 150 km down range &50 km cross range

Mass 400 kg400 kg plus the mass of allreentry and range extension

components

TABLE 3 - AVIONICS SECTION SPECIFICATIONS

Parameter Delivery Order 1 Delivery Order 2Associated Objects No Yes

Size N/A Four 8 cm diameter ortwo 12 cm diameter

Total Mass N/A 40 kg


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Flight termination and active guidance subsystems are driving requirements for cost. TheALFT system should use alternatives where possible that eliminate the need for thesesubsystems.

Each propulsion stage of a given ALFT target will consist of a solid rocket motor. This

proposal will evaluate a variety of existing solid motors as potential solutions to meeteach delivery order. Some of these motors are from current United States governmentstockpiles. These motors are available at zero acquisition cost to the ALFT project andwill be referred to as Government Furnished Equipment (GFE) motors in this proposal.Several other motors evaluated are available for a cost as Commercial Off-The-Shelf(COTS) motors. Table 4 lists the motors evaluated.

TABLE 4 - MOTORS EVALUATED

GFE Motors COTS Motors

Terrier Mk12 OrioleTrident C4 3 rd Stage Castor 1

Improved Orion Castor 4, 4a, 4bATACMS Orion 38

Patriot (PAC-3) Orion 50SR19 Orion 50xlM57 Orion 50sg

Mk11 Mod5 TalosASAT Stage 2 (Altair 3)

To best accomplish the requirements outlined, the ALFT shall consist of a family ofvehicles with each missile satisfying specific missions or launch methods. This shallserve to ensure maximum system capability by not confining a single vehicle with theduty of fulfilling the broad array of missions outlined. Numerous constraints exist to limitand guide the ALFT design. Any solution to the requirements must be in accordance withthe guidelines outlined by applicable documents such as customer supplied SOW andTRD. Also to be considered are applicable range safety documents, environmentregulations, and international treaties.

A detailed breakdown of all of the requirements as well as the teams response to them is provided in Appendix A.


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Identification of Concepts Before attempting to simulate the launch of such a large number of possible missiles, itwas necessary to eliminate combinations that were physically incompatible or thatviolated the requirements of the project. To do this, restrictions were placed on variousaspects of the missile to ensure each option was viable. Then, an interactive tool calledan Interactive Reconfigurable Matrix of Alternatives (IRMA) was created to helpvisualize these constraints and analyze how they affect the size of the design space.Based on this tool, the compatibility constraints could be adjusted until the number of

possibilities was reasonable.

The compatibility constraints placed on the missile were divided into two categories. Thefirst related to the launch vehicle. These restrictions ensured that the rocket would beable to physically launch the required payload and that each stage would be compatiblewith the others. First, any candidate missile needed a thrust to weight ratio of at least 1.2for the first stage. This ensured that the motor provided sufficient thrust to propel the

rocket. The second major restriction limited the ratio of the diameter of two consecutivestages to no less than 0.7 and no greater than 1.4. This constraint ensured that successivestages would fit together without creating excessive drag or necessitating a bulkyinterstage section.

The second category of constraints related to the avionics section and reentry vehicle. Itaccounted for the requirements of each mission as stated in the provided documents. Theavionics section can be separating or non-separating and contain 2 AOs or 4 AOs. Thereentry object can be non-separating or separating, conic or bi-conic. The post-apogeesurvival altitude can either be 150 km, 40 km, or 0 km; this post-apogee survival altitudecan be attained by having an ablative, ceramic, blanket, tile or no TPS system. The range

extension can be possibly be attained by a propulsion system, aerodynamics, or acombination of the two. This propulsion system can be powered by hypergolic

propellants or a compressed gas Reaction Control System (RCS) system. Finally, thetarget range can be 1,000 km, 2,500 km, or maximum range on either a low or hightrajectory.

The filtering on this section of possibilities is directly linked to the requirements. Theavionics section and reentry object do not need to separate for a 1,000 km mission; thismissions avionics section does not carry any associated objects or a maneuveringsystem. The 1,000 km mission specifies a conic reentry object. The 2,500 and maximumrange mission require a separating avionics section with associated objects and a

separating reentry object. This reentry object must have a maneuvering system; thecommittees requirements specified a propulsive reentry system. This reentry objectmust be capable of a post-apogee survival altitude of 40 km, and must have a bi-conicshape.

From these constraints, the IRMA tool was created, as shown in Figure 6 . This toolallowed the user to select one or more possible characteristics of the missile, and then


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display what options are incompatible with this choice and how the design space isaffected.

Figure 7 shows the IRMA with some options chosen, and others eliminated by thecompatibility constraints. The number of design options is reduced significantly. Using

the IRMA, it was possible to adjust the constraints until the size of the list of possibleoptions was reasonable. Then, these options could be run through the simulationenvironment to determine their performance.

FIGURE 6 - IRMA TOOL

FIGURE 7 - IRMA WITH COMPATIBILITY CONSTRAINTS


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Modeling and Simulation Architecture The thrust-to-weight and geometry filtering criteria previously mentioned filters the largedesign space down to 1187 options. The feasibility and performance of these missilesmust then be quantitatively evaluated. This evaluation is conducted using an extensive

physics-based modeling and simulation suite of tools, comprised of seven differentdisciplinary tools linked together in the MATLAB environment. The seven disciplinesfeatured are propulsion, geometry, aerodynamics, trajectory, range extension, thermal,and computer aided design (CAD). The flow of data through this modeling andsimulation (M&S) environment is illustrated in Figure 8 . Each of the seven missiledisciplinary tools will be described in more detail in the following sections.

The most important result of the M&S environment is the maximum range of the givenmissile, which is calculated from the launch trajectory analysis and passed to the trade-offtool. The maximum range determines if the missile is able to meet the delivery orderrequirements and also serves as a discriminator between all of the feasible missiles. In

addition, the reentry trajectory code determines if a given missile can meet the rangeextension requirements. If the range extension can in fact be met for a given missile, theweight of the necessary propellant and thermal protection system (TPS) are computedand also used as discriminators. The reentry trajectory code computes the burn time andin turn the amount propellant used, while the thermal analysis code determines the typeand weight of the TPS. These metrics of range, propellant mass used, and TPS weightalong with other launch vehicle and cost properties are then used to populate aninteractive trade-off tool used in the downselection process.

FIGURE 8 - MODELING AND SIMULATION ARCHITECTURE


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PropulsionThe first element in the modeling and simulation environment is the propulsion code.This software is a database of technical information relating to the nine GFE and tencommercial boosters. This data was obtained from extensive research into public domainsources.

Even after this research, a few details relating to some of the motors that are currentlyactive in the military could not be found. This information was then interpolated fromavailable data on similar systems in order to appropriately characterize these motors.

Motor DatabaseA database of solid rocket motors has also been compiled from previous Georgia Techmissile design teams. The database contained the following pertinent data: gross mass,empty mass, vacuum thrust, specific impulse, burn time, diameter, length, and productionstatus. Although none of the recommended motors were available in this database, someof the motors in the database were chosen to supplement the initial list of motors.Additionally, this information was used to create equations relating different parametersin various motors classes. These fits were then used to fill in the missing parameters forthe recommended motors.

Nozzle Diameter Determination Nozzle exit areas are important in the calculating motor performance. However, they arealso extremely difficult to obtain from public domain sources. As a result, it wasnecessary to calculate these areas from available pictures of the missiles. An example ofthis calculation is illustrated in Figure 9 , which shows the motor and nozzle areas for theATACMS missile. From this picture, the ratio of the nozzle diameter to the motordiameter was approximated. Since the missile diameter was already known to be 0.61 m,the nozzle exit diameter and area could then be calculated.

FIGURE 9 - DETERMINING ATACMS EXIT AREA


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module and initializes the missile with the correct reentry object and motors. At each ofthe subsequent stage separations, the trajectory code recalls the geometry module andcreates a new missile. This is illustrated in Figure 10 .

FIGURE 10 - GEOMETRY CODE OPERATION

Payload Details

The shape of the payloads for both delivery orders is the result of bi-conic reentry objectresearch conducted by the missile design team. A summary of this research was providedto the MSTC shortly after the SOW and TRD documents were received. Included in thissummary was Figure 11 showing a bi-conic reentry object.

FIGURE 11 - NASA BI-CONIC REENTRY OBJECT [9]

The MSTC then specified the shape of the bi-conic reentry object to be a scaled versionof the drawing in Figure 11 with a base diameter of 30 (0.762 meters). The two frustumangles and the tip radius would remain the same. An avionics section of 0.3 meters inlength was appended to the bottom of the reentry object, completing the front sectiongeometry. The bi-conic front section drawing is shown in Figure 12 . This geometry isused for DO2 simulations.


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FIGURE 12 - ALFT BI-CONIC REENTRY OBJECT (DIMENSIONS IN METERS)

For the conic payload, the base diameter (0.762 meters), tip radius (0.057 meters), andtop frustum angle (12.84) were retained from the bi-conic design. The 0.3 meteravionics section was also added. The conic front section drawing is shown in Figure 13 .This geometry is used for DO1 simulations.

FIGURE 13 - ALFT CONIC REENTRY OBJECT (DIMENSIONS IN METERS)

The front section (FS) mass is different for DO1 (conic reentry object) and DO2 (bi-conicreentry object). As noted in Table 1 , the FS mass for DO1 is 400 kg, but the FS mass forDO2 is 400 kg plus the mass of all range extension and reentry survival (heat shield)components. To execute the trajectory simulation, a mass estimate for these componentswas created and is summarized in Table 6 . More detail concerning the composition andcomponent layout of the payloads will be discussed later.


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TABLE 6 - GEOMETRY MASS ESTIMATION

Component Mass [kg]

RO / AS Structure 400Associated Objects 40

Propulsion Subsystems 15Propellant 30

Thermal Protection System 10RO Attitude Control 15

Miscellaneous Electronics 40Total 550

Launch Vehicle and Interstage Assumptions

All motors and payloads are assumed to have a constant density. Each of the interstagesis assumed to have a constant length of 0.2 m and a constant density of 1500 kg/m 3 [10].The relatively short length of the interstages is used assuming that the nozzle isincorporated into the researched or provided motor lengths. Therefore, the interstageonly needs to be long enough to provide room for the separating mechanism. Also, thisinterstage length worked consistently with the MDATCOM aerodynamic code whileother interstage lengths generated persistent errors.

Aerodynamics The aerodynamics analysis consisted of determining the coefficient of drag (C D) of themissile as the configuration changed from the full missile at launch, through each of thestage separations, and finally to the configuration with just the reentry object. For eachof the missile configurations, a table of C D values corresponding to a range of Machnumbers and altitudes was created. A plot of a notional C D table is shown in Figure 14 .The red line indicates the missiles corresponding trajectory through this flight regime.Because the missile is symmetric about its central axis, and the flight path angle isassumed to be aligned with the axis of the missile, there was no need to calculate acoefficient of lift or aerodynamic moments.

Missile DATCOMMDATCOM is a flexible aerodynamic code that enables quick evaluation of simplegeometries. The shape of an axisymmetric body can be defined by a list of points (asdescribed in the next section), which allows the missile geometry to be varied to matchall the combinations of boosters being evaluated. It is a legacy compiled code, though,which can lead to difficulties when certain missile shapes cause it to fail for unknownreasons. If the failures were isolated, the points were interpolated from the surroundingvalues in the drag table. But when a given missile shape failed for all Mach and altitudecombinations, the drag table was replaced by that of a similarly-sized missiles table.


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FIGURE 14 - DRAG COEFFICIENT THROUGHOUT THE FLIGHT REGIME

Missile ShapeWithin the programming confines of MDATCOM, there are several ways to define the

shape of a missile. The most reliable method found for the purposes of this study was tosupply a set of points defining the longitudinal (X) and radial (R) coordinates. The shapeof the RO was constant, but the shape of the missile changed for each combination andfor each stage of the flight. The X and R data for the missile shape was produced by thegeometry code as previously explained. This data was then passed into the MATLABwrapper for MDATCOM, which then attached the RO shape to the missile body shape.Figure 15 shows how these original points were expanded to best define the entire missileshape. The original data points are the green circles (only points of inflection weregiven). Then intermediate points were interpolated, starting at the center of the longestspan. The interpolated points are shown as red Xs (MDATCOM input is limited to 50

points). Internally, MDATCOM would draw the shape in more detail from the points

given. Those points are connected by the blue line.

FIGURE 15 - MISSILE DATCOM SHAPE DEFINITION


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Trajectory The trajectory module of the simulation integrated the equations of motion of the missileto produce a complete trajectory. It accepted the stages comprising the missile as inputs,as well as the payload the missile must lift. From this data, it called the geometry and

propulsion modules, which provided the necessary weight and thrust data for each stage,and then called the aerodynamics module, which provided a drag model for each stage.In addition, the longitude and latitude of the launch site and the launch angles, azimuthand elevation, were provided as additional inputs. The propagation of the trajectory wasthen divided into three phases. The first phase, or the boost phase, simulated the launchof the rocket and integrated the trajectory until either the rocket reached 100 km altitudeor the burnout of the final stage, whichever came later. The second phase, or midcourseand reentry phase then propagated the missiles motion until it reached zero altitude. Thefinal phase, or the range extension phase, simulated a burn performed beginning atapogee to attempt to reach the cross range and down range extensions required for theALFT system. Each of the three phases is described in further detail below.

The trajectories for DO1, DO2, and maximum range for a notional missile are shownusing Google Earth in Figure 16 .

FIGURE 16 - 1000 KM (RED), 2500 KM (YELLOW), AND MAX RANGE (GREEN) TRAJECTORIESFOR A NOTIONAL MISSILE

This module of the simulation was written in MATLAB like many of the other modelingtools. Various components were common to all three phases. The integrator used afourth order, variable step, Runge-Kutta method, which provided for both excellent


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numerical accuracy and reasonable speed. Also, all phases made use of an oblatespheroid model of the Earth, which assumed an equatorial radius of 6,378.145 km and a

polar radius of 6,356.785 km.

Boost Phase

The first phase of the simulation models the boost segment. Since the ALFT system isintended to be unguided, it was assumed that the rockets angle of attack would always bezero; therefore lift could be neglected. Also, jettisoning each stage after it burns outcauses air turbulence which affects the rocket. However, this effect is relativelynegligible and very difficult to model, and so was neglected. In this way, the only effectof jettisoning a spent stage is an instantaneous change in the vehicles mass. Since therocket is initially rotating with the Earth, and the atmospheric forces are calculatedrelative to the planets surface, this phase was integrated in a topocentric frame centeredat the launch site, with the equations of motion given by 305HEquation 1. The missile was to

be launched from a rigid rail, so the rate of change of the flight path angle was fixed atzero for a short period of time after ignition. At each point in the trajectory, the thrust

and drag were calculated based on the data provided by the propulsion and aerodynamicsmodules respectively. These equations were propagated until the missile reached theedge of the atmosphere, defined as 100 km above mean sea level. However, if the rocketwas still burning at this altitude, the equations were propagated until burnout andseparation (if applicable) of the final stage. All position and velocity information wasthen converted into a non-rotating geocentric coordinate frame.

EQUATION 1 - BOOST PHASE EQUATIONS OF MOTION

( ) sincos 2r m D

mT

v T +=&

sinvr &=

( )

+++= sincos1cos 2r m L

mT

vr v

T &

cosr v=&

sp I g T

m0

=&

State Variablesv: velocity magnituder : radius: flight path angle (angle between velocity andhorizontal)

: range angle (angle between line from center of planetto rocket and similar line to launch site)m: mass

Other VariablesT : thrust

D : drag L: lift (assumed zero) : Earth gravitational parameter (3.986x10 5 km3/s2)

: angle of attack (assume zero)T : thrust vector angle (assumed zero) g 0: acceleration of gravity at Earths surface (9.81 m/s

2) I sp: Impulse of motor


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Midcourse and Reentry PhasesThe midcourse and reentry segments of the missiles flight were both modeled by thesecond phase of the trajectory module. These segments were combined into a single

phase because the reentry vehicle possessed no aerodynamic maneuvering systems of its

own306H

[11]. Due to this and the high velocity of reentry, the effect of the atmosphere onthe trajectory and range of the missile was small. To further simplify the integration,therefore, aerodynamic forces were neglected once the rocket ended the boost phase.While this also implied ignoring the effects of heating upon reentry, these forces areactually negligible. The equations of motion for this phase, given in 307HEquation 2, aretherefore based on Keplers laws of orbital mechanics. These equations were propagateduntil the missile reached zero altitude.

EQUATION 2 - MIDCOURSE EQUATIONS OF MOTION

r r

r v&&v

3=

Range Extension PhaseThe requirements of the ALFT dictated that the missile must be able to obtain both adown range and cross range extension, so the third and final phase of the simulationmodeled a deviation from the trajectory integrated in the first two phases using amaneuvering thruster on the reentry vehicle. The data for the thruster was provided bythe propulsion module. By beginning the burn at apogee of the previously calculatedtrajectory, this phase of the simulation calculated the required burn time and mass of fuelto achieve the down range or cross range extension. In addition, it also calculated therequirements for achieving both simultaneously. The assumptions and equations ofmotion for this phase were similar to those for the second phase, with the exception that

thrust was added to the equations of motion, provided in 308HEquation 3.

EQUATION 3 - RANGE EXTENSION EQUATIONS OF MOTION

r r m

T r

vv

&&v

3

=

An illustration of the range extension capabilities for a notional missile is shown in309HFigure 17.


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FIGURE 17 - DOWNRANGE (RED), CROSSRANGE (BLUE), BOTH (PURPLE), AND ORIGINAL

(GREEN) TRAJECTORIES FOR A NOTIONAL MISSILEUpon completion of all three phases of the simulation, the range and apogee of themissile as well as the requirements for each of the three range extensions were returned.In addition, other intermediate calculations, such as the position and velocity, the forcesof the missile, and the mass of the missile as function of time throughout the trajectory,were provided. Since the project requirements specified meeting a certain range, anoptimizer was used with the trajectory module to find the appropriate launch elevationsthat would result in both the target range and the maximum range.

Code EvaluationTo determine if the trajectory code performed as expected, it was evaluated using knowntrajectory data. This data came from the NASA Sounding Rocket Handbook thatincludes performance graphs for several sounding rockets [11] . The single stage BlackBrant VC MK1 data included all of the physical characteristics of the motor: length,diameter, gross weight, propellant weight, average thrust, and burn time. The

performance graph included impact range, apogee altitude, and time above 100km for anarray of points varying the payload weight and launch angle.

The physical characteristics of the Black Brant VC MK1 were added to the motordatabase, and then the trajectory code was iterated through all of the payloadweight/launch angle points. The calculated trajectory data (impact range and apogeealtitude) were compared to the data in the Sounding Rocket Handbook and are also

plotted in Figure 18 [11] .


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FIGURE 18 - CODE RESULTS VS. KNOWN DATA FOR BLACK BRANT VC MK1

Overall, the trajectory code does very well at reproducing results consistent with knowndata. The worst case data point is at maximum payload mass (589.7kg) and lowestlaunch angle (76). The results for this point deviate 5.4% for range and 10.1% forapogee.

Thermal Analysis One of the most difficult phases of ballistic flight is atmospheric reentry. An objectreentering Earths atmosphere begins to experience drag and aerothermal effects belowapproximately 100 km, where the atmosphere begins to thicken to a substantial density atfor an object traveling at high velocities. The temperature in the boundary layer aroundthe reentering object can reach very high values, up to tens of thousands of degreesFahrenheit, and because there is high surface shear, there is significant heat transfer to theRO surface [13] . In order for the reentering object to survive to a desired altitude with its

payload in tact, this heat transfer must be controlled to keep the RO structuraltemperature lower then its failure point, and to keep the internal temperature lower thenthe payloads failure point.

There are many forms of protection that can be employed, collectively referred to as TPS.For a ballistic reentry, such as is employed for the ALFT, the two most applicable formsof TPS are passive and active systems.

Active systems in general consist of some sort of insulating material that is integratedwith a system that keeps that insulating material below a certain temperature. Suchcooling systems include refrigeration, cryogenic fluids, and flash evaporation. Activesystems have the benefit that they can be used in extreme reentry cases where insulating


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material alone is not sufficient to withstand the heat load and maximum heat flux. Thedownfall of active systems is that they are complex and heavy because of the need forstorage of refrigerant fluids, as well as the movement of fluids through the insulatingmaterial. This adds cost, weight, and uncertainty of success.

Passive systems consist of some sort of insulating or ablative material. Passive systemsabsorb and remove heat from the RO structure without the benefit of moving parts oradditional systems, and so are desirable when possible because of weight and complexitysavings. For this reason, passive systems, using either thermal soak or ablativetechnologies, were chosen for the TPS of the ALFT.

Assuming the RO structure is strong enough to survive the thermal environment duringlaunch, for the non-separating conic RO DO, aerothermal heating is not an importantissue because it is only required to achieve a survival altitude of 150 km, and so nothermal analysis was done for the conic RO. For the separating bi-conic RO deliveryorder however, aerothermal heating must be taken into account in order for the RO to

reach the desired survival altitude of 40 km.Thermal analysis of the RO is one of the last steps in the ALFT M&S environment,taking in the results from the reentry trajectory and geometry codes in order to analyzethe reentry thermal environment and the type and amount of TPS required to allow theRO to survive to the desired survival altitude. The design tool selected for TPS designand sizing is NASAs MINIVER, an aerothermal analysis and conceptual design tool. Inaddition, a zero-order conceptual analysis was written to supplement this analysis. Bothanalyses used a common TPS database from which to size and select.

Thermal Protection System DatabaseThe TPS database used for the TPS sizing was developed using NASAs TPSX MaterialProperties Database, Web Edition Version 4, which was developed by the ThermalProtection Materials & Systems Branch at NASA Ames Research Center, along with

NASA Langley Research Center, and the NASA Office of the Chief Engineer. It consistsof an extensive list of TPS types and properties, drawing from the NASA Ames ThermalProtection Materials database and the NASA Johnson Space Center PathFinder Materialsdatabase. In order to be of use in the analysis, the density, specific heat, thermalconductivity, emissivity, and single use maximum temperature limit of each materialneeded to be available. From the TPSX database, 49 materials had the necessaryinformation.

Figure 19 shows a screenshot of the assembled database. It should be noted that thisdatabase also includes ablative materials. The single use maximum temperature given forthese materials is actually the temperature at which the material begins to ablate. For thedegree of fidelity required for the conceptual design of the ALFT, this use of the ablationtemperature actually treats ablative materials as pure insulators. If the analysis selects anablative material, it will then oversize the ablative material, but not grossly so. The resultshould be on the order of magnitude of a more detailed design analysis.


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FIGURE 19 - TPS DATABASE CREATED FOR SIZING AND ANALYSIS BASED UPON THE NASATPSX MATERIAL PROPERTIES DATABASE

Zero-Order Conceptual AnalysisThe zero-order thermal analysis tool was written in MATLAB and was based heavily onthe theory and assumptions represented in references [13][14][15] . It is a one-dimensional analysis that sizes TPS based on material type, surface temperature of theTPS material, the maximum temperature allowable at the RO structure surface, radius of

the RO nose, velocity, and density at the corresponding altitude.

Key AssumptionsSeveral key assumptions are used to simplify the analysis. The aerothermal effectsexperienced by the RO during launch are taken to be negligible compared to the heatingthat the RO is subjected to during reentry. Also, the nose is assumed to experience thehighest temperature. To simplify the calculation of the weight of the TPS, it assumed thatthe TPS thickness is constant over the reentry surface of the RO, which is equal to thethickness of the TPS required to withstand the heat input at the RO nose. The TPS massis assumed to absorb heat uniformly. The atmospheric density was calculated using thestandard atmosphere model.

Solution ProcedureThe maximum temperature experienced by the RO is calculated using the recoverytemperature relation described in reference [16] . Using this temperature, the maximumtemperature allowable for the RO structure, and the conductivity of each TPS type, themaximum thickness necessary for each TPS type is calculated. Using that thickness, theexposed area, and the density of each TPS type, mass is found. The TPS typecorresponding to the lowest mass is then selected.

Higher Fidelity AnalysisMINIVER was used in conjunction with the ASDLs Thermal Management System Sizer

(TMSS) for the higher fidelity analysis. It has been used in the past for both governmentand industry projects, and has given results that agree with more detailed solutions for

projects such as Space shuttle, HL-20, X33, X34, X37, and X43 [17] . TMSS is aMATLAB code written and developed by Irian Ordaz of ASDL that generates theMINIVER input geometry and uses the MINIVER outputs to size and select TPS typeusing the 1D transient heat equation.


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First the RO geometry was constructed using Visual Sketch Pad (VSP) and the mission profile was created using the reentry optimization code by making a table of time,altitude, and velocity, ending where altitude is equal to 40 km. TMSS then generatesstreamline information from the VSP geometry. MINIVER uses this streamlineinformation along with the reentry profile to provide aerothermal information for the

entire mission profile. TMSS then uses the MINIVER output to size and select the bestTPS type from the TPS database. This process is illustrated pictorially in Figure 20 .

FIGURE 20 - THERMAL M&S FLOW FROM VSP GEOMETRY (A) TO TMSS STREAMLINEGEOMETRY (B) TO MINIVER THERMAL ANALYSIS (C) AND TMSS TPS SIZING (D)

CAD For this study, Dassault Systemess Computer Aided Three Dimensional InteractiveApplication (CATIA) was used to construct the CAD models of each candidate missiles.CATIA has a built-in tool that allows it to read data from Microsoft Excel spreadsheetsand assign variables to any parameter within the model. This feature enabled rapidupdating of the model to fit the specifications laid out by the most recent run. As a result,CATIA could then parametrically update the model automatically in a matter of seconds.Screenshots of each model were taken for use in the MADM tool and used to help selectthe ALFT systems.

These CAD models allowed the team to visualize the design space of candidate missiles.This visualization served as an aid in the selection process. A snapshot of a small section

A B

C D


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of the design space is shown in Figure 21 . Since there are assumptions in the first sixM&S tools, it is possible that missiles with structural or aeroelastic problems couldappear to perform well. By visualizing the options, these problems could quickly beeliminated. Additionally, CATIA served as a debugging tool for the M&S environment.If a large number of cases failed in one of the other tools, visual examination could

usually help identify the problem. For example, extreme differences in motor radii fromone stage to the next sometimes caused problems for the aerodynamic analysis.

FIGURE 21 - VISUALIZATION OF DESIGN SPACE

In the detailed design phase, CATIA was also used to model the internal layout of themissile FS. The purpose of this coarse model was to determine how to best organize therequired subsystems inside of the RO and AS. Such items included the associated objectsdetailed in the TRD, a propulsive motor, a separating avionics section, sensory andtelemetry gear, and a gyroscope for stability.


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Concept SelectionAfter obtaining all of the results from the modeling and simulation environment, asignificant amount of data was compiled. This data was then analyzed and put through aMADM tool. This is a tool that allows for the judgment of all the possible concepts(note: possible here means that the concepts meet the requirements for the deliveryorders- those cases that cannot meet the range requirements are discarded in a first passfilter). The judgment is parametric because of the influence of user-inputted weightingscenarios. The tool and its architecture are described below, followed by the finalconcepts chosen for further analysis.

Interactive Trade-off Tool In order to judge all of the concepts on their merits, an environment needs to be created inwhich all of the important characteristics of the booster combinations are considered.Such an environment has been created in Microsoft Excel using a combination ofworksheet functions and VBA. The dashboard is shown below in Figure 22 .

FIGURE 22 - PARAMETRIC TRADE TOOL DASHBOARD

Technique for Ordered Preference by Similarity to Ideal SolutionIn order to perform the multi-attribute calculations needed for this kind of evaluation, a

process called Technique for Ordered Preference by Similarity to Ideal Solution(TOPSIS) was used. TOPSIS is a MADM technique that takes the data from its initialform and normalizes and weights the data. Then, it picks out the positive and negativeideals and calculates the distances from these ideals for each alternative. Finally, the rankis based on which alternative is closest to the positive ideal and farthest from the negativeideal. The concept behind TOPSIS (shown only in 2-D here, but is actually multi-


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dimensional) is shown here in Figure 23 . The figure shows the concept of taking theconcept with the farthest distance from the negative ideal and the closest to the positiveideal.

FIGURE 23 - PARETO FRONTIER

Thus, using TOPSIS, a ranked order of solutions can be presented for the individualdelivery order missions. However, an additional calculation regarding commonality ofthe boosters can be used, as is discussed later.

Weighting ScenariosOne of the ways in which this tool will help is that the weighting scenarios will be able to

be changed on the fly and results instantly used. Using integrated macros and code, thetool will run anytime anything is changed on the sheet, thus creating an environment inwhich trade-offs can be discussed in real time.

Essential to the form and function of TOPSIS are the weighting scenarios that are to beimplemented to each criterion at hand. The weighting scenario part of the tool is shown

below in Figure 24 .

FIGURE 24 - WEIGHTING SCENARIOS


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Using the weightings shown on the right, TOPSIS runs on the fly and calculates a newand different set of rankings for each delivery order.

Delivery Order CommonalityAfter TOPSIS is performed, ranked lists of importance are given for each of the delivery

order specifications. However, as the title of this project contains the importantcharacteristic of being affordable, it is necessary to consider the common boosters thatare ranked highly in the TOPSIS results, to see if it is possible to stack boosters or usethe same boosters to perform all of the delivery order missions. This is done bycalculating all of the possible common boosters in a given data-set using MATLAB, andthen taking the rankings of each from their TOPSIS result, and adding them together, thusgiving the best family of rocket boosters overall. This is a useful but not necessary stepin picking out the final down-selection result.

Architecture of the ToolFigure 25 shows how the actual tool works to output the results shown on the dashboard.

In essence, the data from the first and second delivery order missions are taken into thespreadsheet and run through TOPSIS. Then, the results for the pre-determined familiesof rockets are assigned and re-sorted. This will yield a final concept list from which onecan pick a family of rockets that both fits the delivery order criteria and is affordable.

FIGURE 25 - CONCEPT SELECTION FLOWCHART

Missile DownselectionIn the downselection process, several factors had to be weighted. Low cost requirementsas well as transportability requirements were able to be weighed against the performancerequirements.

First the team only looked at the transportability and reliability by only looking atminimum stage vehicles. The Orion50sg and the Castor 4 series of motors can fulfillDO1 requirements as a single stage. Similarly, these motors will be a first stage for a


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first stage and either an SR19 or M57 second stage as well as an SR19 first stage and anM57 second stage can fulfill DO1 requirements. Other three-stage combinations of GFEwork for DO1, but these were decided against because the TRD requires a two-stagemissile for DO1. Similarly, SR19 and M57 based missiles are capable of meeting DO2mission requirements.

FIGURE 26 - PROOF OF FUNCTIONALITY

Also, because a driving requirement is cost, it is imperative to look at the solutions withthe best cost-factor, so in a similar fashion it is possible to look at the solutions with theleast amount of cost associated with them, shown in Figure 27 .

FIGURE 27 - BEST COST OPTIONS

After playing with several different weighting scenarios, two concepts for DO1 areconsistently at the top of the rankings: a Talos first stage, M57 second stage and a SR19first stage and an M57 second stage. Figure 28 and Figure 29 show these three motors.


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FIGURE 28 - TALOS SOLID ROCKET MOTOR [18]

FIGURE 29 - SR19 AND M57 SOLID ROCKET MOTORS [18] [19]

Since both concepts are two stage missiles made out of GFE, the decision came down todiscriminators in the operations and transportability of these two concepts. TheSR19/M57 missile is a much more powerful concept with a predicted maximum range ofover 2,400 km. This concept flies a maximum range trajectory at an initial launch


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elevation of 85.9, and the targeted 1,000 km mission flies at an initial launch elevationof 88.8. A variation in range of 1,400 km in less than 3 of initial elevation means thatthe trajectory is extremely sensitive to an initial launch elevation.

In contrast, the Talos/M57 concept has a maximum range of 1,294 km at a launch

elevation of 64.0 and a targeted launch elevation of 74.6. The Talos/M57 concept has amuch lower sensitivity to deviations of initial launch elevation which makes it a superiorchoice given the requirement that this will need to be launched from the deck of a rollingship. The data for this is shown in Table 9 .

TABLE 9 - LAUNCH ANGLE SENSITIVITY (DO1)

Motor Configuration[First Stage /

Second Stage]

Maximum Range[km]

Launch Angle forMaximum Range

[degrees]

Launch AngleTargeted to

1000km [degrees]

Talos / M57 1294 64.0 74.6SR19 / M57 2475 85.9 88.8

In addition to the favorable launch operations, the Talos/M57 is 5,174 kg less massiveand 0.77 m shorter than the SR19/M57 making the Talos/M57 easier to transport, asshown in Table 10 . The Talos/M57 concepts superior non-performance characteristicsmake it the best concept for DO1. The CAD for this concept is shown in Figure 30 .

TABLE 10 - TRANSPORTABILITY COMPARISON (DO1)

Motor Configuration[First Stage / Second Stage]

Total Launch Length[m]

Gross Liftoff Weight[kg]

Talos / M57 8.23 4738SR19 / M57 9.00 9912

When considering DO2, two GFE based concepts are consistently at the top of theTOPSIS results: Talos/SR19/M57 and Talos/SR19/SR19. Both missiles haveapproximately the same range so the main discriminators came down to operations andtransportability. The Talos/SR19/M57 has a maximum range trajectory launch elevationof 75.2 and a 2,500 km targeted trajectory launch elevation of 83.3. TheTalos/SR19/M57 has a maximum range trajectory launch elevation of 84.9 and atargeted trajectory launch elevation of 87.7. The Talos/SR19/M57 option is lesssensitive to launch elevation angle deviations than the Talos/SR19/SR19, as shown in

Table 11 .


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FIGURE 30 - TALOS/M57, DO1 CONCEPT

TABLE 11 - LAUNCH ANGLE SENSITIVITY (DO2)

Motor Configuration[First Stage / Second Stage /

Third Stage]MaximumRange [kg]

Launch Angle

for MaximumRange[degrees]

Launch Angle

Targeted to 2500km(High Trajectory)[degrees]

Talos / SR19 / M57 3632 75.2 83.3Talos / SR19 / SR19 3617 84.9 87.7

The Talos/SR19/M57 is also 5,000 kg less massive and 1.42 meters shorter than theTalos/SR19/SR19 concept, as shown in Table 12 .

TABLE 12 - TRANSPORTABILITY COMPARISON (DO2)

Motor Configuration[First Stage / Second Stage /

Third Stage]Total Launch Length[m] Gross Liftoff Weight[kg]

Talos / SR19 / M57 12.90 12322Talos / SR19 / SR19 14.32 17555

The resulting RO characteristics were also looked at while making a decision, as showninTable 13 . However, these characteristics ended up being a non-discriminator because thedifference in required fuel is less than a kilogram and the difference in required TPS isless than a tenth of a kilogram.

TABLE 13 - RO CHARACTERISTICS COMPARISON

Motor ConfigurationFirst Stage / Second Stage /

Third Stage]

Range ExtensionMotor Propellant Weight [kg] TPS Weight [kg]

Talos / SR19 / M57 39.7 11.0Talos / SR19 / SR19 40.5 11.1


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The Talos/SR19/M57 concepts strengths in targeting and transportability as well as theM57 to avionics section interstage commonality make it the best concept for DO2. It isshown in CAD in Figure 31 .

FIGURE 31 - TALOS/SR19/M57, DO2 CONCEPT

As a final sense check, it was necessary to make sure that using the relatively small Talos booster with larger stages above it would actually work. In order to see if this was possible, research was done to see if other configurations in use today or in the past haveused the Talos as a primary booster. Several examples were found to show that the Talos

booster was used to propel multiple large stages and payloads above it. One example isthe Starbird, a four stage configuration of: Talos / Sargent / Orbus1 / Orbus1 / Payload.A picture of it is shown in Figure 32 .

FIGURE 32 - STARBIRD, ILLUSTRATING TALOS AS FIRST STAGE


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Missile Conceptual Design

Interstage Design Now that a motor configuration for each delivery order has been selected, some designdetails for the missile can be determined. A more accurate estimate of the interstageweights is calculated based upon improved geometry and preliminary structural analysis.

The maximum half-angle for each interstage is 16.7. This value is chosen based uponhistorical research of Titan and Minuteman missiles performed by the 2006-2007 GeorgiaTech missile design team. For configurations where the interstage is connecting twocomponents of similar diameter, this value is too large which results in a very shortinterstage. For these cases the half-angle is reduced until the minimum interstage lengthis 0.5 meters. Table 14 and Table 15 show the final geometry for each interstage forDO1 and DO2 respectively.

TABLE 14 - FINAL GEOMETRY FOR INTERSTAGES (DO1)Interstage Location Half-angle [degrees] Length [m]

Payload M57 13.3 0.5M57-Talos 11.8 0.5

TABLE 15 - FINAL GEOMETRY FOR INTERSTAGES (DO2)

Interstage Location Half-angle [degrees] Length [m]Payload M57 13.3 0.5

M57 SR19 16.7 0.55SR-19 - Talos 16.7 0.9

Two stress calculations are performed to determine the acceptable wall thickness of theinterstage frustums. The first calculation determines the compressive stress at maximumdynamic pressure (max Q) for each interstage. This calculation is based upon the dragforce and acceleration which are determined from the modeling and simulation results.Because the selected configuration for DO1 and DO2 both use the Talos first stage, maxQ occurs at the same point for each flight profile. This point is six seconds into the flightat the end of the Talos burn.

The second calculation takes into account the maximum allowable buckling stress due toaxial compression. This calculation is based upon the diameter to thickness ratio and the

mechanical properties of the material selected. For each calculation, 4130 steel was theselected material which is commonly used for solid rocket motor casings and interstages.

For each interstage design the maximum allowable buckling stress is the driver indetermining the minimum interstage thickness. Once this minimum thickness isdetermined, the frustum shell thickness is determined using the previously determinedgeometry and the density of 4130 steel. This value is then multiplied by 1.3 to accountfor the structure of the mating surfa

affordable low fidelity target design competition final report

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