AIAA Missile Systems Technical Committee (MSTC) … Documents...A Proposal in Response to AIAA Missile Systems Technical Committee (MSTC) 2007–2008 Graduate Missile Design Competition

A Proposal in Response to

AIAA Missile Systems Technical Committee (MSTC) 2007–2008 Graduate Missile Design Competition

Affordable Low Fidelity Target Systems (ALFT)

AEROSPACE SYSTEMS DESIGN LABORATORY

SCHOOL OF AEROSPACE ENGINEERING GEORGIA INSTITUTE OF TECHNOLOGY

ATLANTA, GA 30332-0150

1 JUNE 2008

2007/2008 AIAA MSTC ALFT Missile Design Competition

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Executive Summary Target 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 for the evaluation of the BMD system’s performance at a level that is beyond the capabilities of advanced computer modeling and more representative of an actual engagement. Drawbacks for testing with target missiles include the high costs incurred and the difficulty of using targets to simulate a large variety of systems. Therefore, a new target system that is both affordable and flexible is desired. The proposed Affordable Low Fidelity 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 family of ALFT systems in the AIAA/Missile Systems Technical Committee (MSTC) Graduate Team Missile Design Competition during the 2007-2008 academic year that addresses these perceived target drawbacks. In order to conceptually design this target family an appropriate process was developed. First, in-depth research into the problem, including motor and payload front section characteristics, was compiled from the open literature. A Quality Function Deployment (QFD) was also used to map the requirements to the significant engineering characteristics. This information allowed the team to generate a large design space of vehicle alternatives. These options were mapped and a morphological analysis was conducted to identify the compatible options. The feasible designs were then evaluated using an extensive modeling and simulation environment that appropriately addressed the physics of all relevant disciplines for the target family’s performance. After this analysis, a family of vehicles was downselected from the feasible designs using multi-attribute decision making techniques. A higher fidelity design was then carried out on the most promising families of targets. Special attention was paid to the 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 sounding rocket equipment to meet the needs of an assortment of missions because of their low cost and asset availability. To further reduce the target family cost, the targets were designed to be unguided. This deviation from most targets currently in use eliminates the need for a complex and costly active guidance system. This study focused on two delivery orders (DOs): the first (DO1) requiring a minimum range of 1000 km, and the second (DO2) requiring a minimum range of 2500 km. Both DOs featured payloads greater than 400 kg. These payloads included a reentry object (RO) and an avionics section (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 common first and final stage for each. The solution to DO1 consists of a Talos first stage and an M57 second stage, with a maximum range of 1488 km. The solution to DO2 consists of a Talos first stage, SR19 second stage, and M57 third stage, with a maximum range of 4166.


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Table of Contents Executive 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 Figures Figure 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 Tables Table 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 Acronyms ALFT Affordable Low Fidelity Target ASDL Aerospace Systems Design Laboratory AO Associated Object AS Avionics Section BMD Ballistic Missile Defense CAD Computer Aided Design CATIA Computer Aided Three Dimensional Interactive Application CFP Contractor Furnished Property CONOPS Concept of Operations COTS Commercial Off-The-Shelf DO Delivery Order DoD Department of Defense EAFB Eglin AFB EMC Electromagnetic Compatibility EMI Electromagnetic Interface ES Experimental Subsystems EWR Eastern and Western Range FS Front Section GFE Government Furnished Equipment GFP Government Furnished Property HILM Hit Impact Location Measurement IRMA Integrated Reconfigurable Matrix of Alternatives M&S Modeling and Simulation MADM Multi-Attribute Decision Making MDA Missile Defense Agency MDATCOM Missile Data Compendium MMH Monomethyl Hydrazine MON Mixed Oxides of Nitrogen MSTC Missile Systems Technical Committee OTS Off-The-Shelf PAC-3 Patriot Advanced Capability-3 PMRF Pacific Missile Range Facility QFD Quality Function Deployment RCS Reaction Control System RO Reentry Object RTS Reagan Test Site SE Support Equipment SOW Statement of Work TMSS Thermal Management System Sizer TOPSIS Technique for Order Preference by Similarity to Ideal Solution TPS Thermal Protection System TRD Technical Requirements Document V&V Verification and Validation VAFB Vandenberg Air Force Base


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VLDE Very Low Density Elastomeric VSP Visual Sketch Pad WFF NASA/Wallops Flight Facility WMD Weapon of Mass Destruction WSMR White Sands Missile Range

2007/2008 AIAA MSTC ALFT Graduate Missile Design Competition

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Conceptual Design Team The following proposal summarizes the work performed for the 2007-2008 AIAA MSTC Missile Graduate Design Competition. The conceptual design team for this competition consisted of graduate students and undergraduate students from the School of Aerospace Engineering at the Georgia Institute of Technology. Combined, these students contributed more 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 Mavris Professor and Boeing Professor of Advanced Aerospace Systems Director, Aerospace Systems Design Laboratory (ASDL) Georgia Institute of Technology

ALFT Design Team Members and Responsibilities Mr. Adam Maser Program Manager Mr. Billy Gallagher Chief Engineer Mr. Frank Coleman Propulsion Mr. Lee Demory Structures Mr. Andrew Hensley Propulsion Mr. Andrew Herron Thermal Analysis Mr. Kamal Kayat Aerodynamics Mr. Brad Robertson Trajectory Ms. Elizabeth Saltmarsh* Preliminary Design Mr. Doug Stranghoener* Visualization Mr. Rob Willett* CAD

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

*Denotes Undergraduate Team Member


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Introduction The following proposal includes engineering analysis and hardware design associated with the buildup and launch of flight vehicles in support of ALFT missions. The design solutions were created in accordance with the Statement of Work (SOW), Technical Requirements Document (TRD), December Kickoff Meeting discussions, and March Systems Requirements Review discussions. Concepts were explored with the knowledge that 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/Wallops Flight 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 capabilities were also explored. Also, because flight termination and active guidance sub-systems are driving requirements for cost, alternative solutions were explored to eliminate the need for these sub-systems.

ALFT Overview The ALFT system consists of a family of affordable non-separating and separating target vehicles designed to complete two representative missions from 1000 to 2500 km in range always using ground launch techniques, but also possessing air and sea launch capabilities if possible. The payload of the ALFT system includes a maneuverable and range extension capability RO and AS, with the ability to deploy AOs. The objective of the ALFT system is to provide a low cost, quick turn-around missile system that can be used for assessing and calibrating sensor system developments and modifications, payload developments, sounding rocket experiments, and limited intercept experiments.

The Need for Missile Defense The current ballistic missile concern is different from the concern prevalent during the Cold War. Wartime enmity causes state leadership to be more risk prone, and unstable governments can result in potential change in control of the military forces. Weapons of mass 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 so WMDs are used to compensate for conventional strength. WMDs can also be used as a means to coerce the United States and its allies. Missile defense serves as an enabler of United 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 sense in a multipolar world of proliferating nuclear powers. Mutual destruction is not likely to work against religious fanatics; desperate leaders may blackmail with nuclear weapons; blackmail or accidents could run out of control. And when these dangers materialize, the refusal 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 rapid expansion 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 missile capability has spread throughout the world, from just a few nations to a major percentage of 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 HAVE

BALLISTIC MISSILE RANGES >1000KM [1] The greatest strategic threat to the United States is an attack by one or more ballistic missiles armed with nuclear or other weapons of mass destruction. Today, the United States remains vulnerable to this form of attack. Thus there is an urgent need for robust and layered missile defenses. Systems based on land, sea, air, and in space which are capable of intercepting a missile during any phase of its flight are necessary to establish a reliable defense. As the United States’ Missile Defense Agency (MDA) advances with development and deployment of its ballistic missile defense systems, a need is created to test and evaluate these fast emerging systems. With interceptor missile capabilities proceeding at a rapid pace, an ALFT system is needed for real threat simulation. There are two prominent forms of testing a missile defense system: simulations and computational analysis, and using physical targets.


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FIGURE 2 - WORLDWIDE BALLISTIC MISSILE CAPABILITIES AS OF 2005, WHERE NATIONS IN 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 make many assumptions and use theoretical models, both of which introduce uncertainty. Physical targets are beneficial because they truly evaluate a system’s performance in reality. 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 a computational simulation and a physical target test. Essentially, simulations are ideal for sizing and selection of preliminary designs, but a physical target will always be needed to validate a BMD system.

FIGURE 3 - LEFT: BMD COMPUTATIONAL SIMULATION USED BY THE ASDL; RIGHT: IMAGE

OF 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 the motors. 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 target missiles 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 most common uses are as targets, experiment delivery vehicles and sensor systems test cuing objects. The objective of the ALFT program is to provide low cost, quick turn-around missile systems that can be used for assessing and calibrating sensor system developments and modifications, payload developments, sounding rocket experiments, and limited intercept experiments. Hence a new target system will play a critical role in safeguarding the United States and its allies in the twenty-first century.

Request for Proposal The 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. Due to the rapidly advancing progress and needs of the United States’ missile defense operations, the ALFT shall be a flexible, cost efficient and quick turn-around missile systems. The capability of launch from non-terrestrial platforms, which may be executed via 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 and government furnished properties, such as WSMR and VAFB, for more traditional launch operations. Regardless of launch mode, the ALFT shall be ready for launch within 20 days of call-up up from the long-term storage condition. The system shall have a calculated 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 Methodology An appropriate process was developed to design the ALFT systems. This process was divided into three distinct phases, each of which concluded with an industry review. Phase 0 was the problem definition phase, which involved the background research into BMD and targets of interest and the definition of the two requirements documents: the SOW and the TRD. Phase I was the concept selection phase. During this phase, modeling and 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 missile front section.

Problem Definition At the beginning of the design process, the team began by conducting in-depth research into BMD systems and current target vehicles. This allowed the team to identify the shortfalls of current systems and the challenged faced in the design of the ALFT systems. After the background research was completed, the requirements documents, the SOW and 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 also developed during this time. Due to the high importance that the requirements place on the design of the reentry object and the classified nature of reentry object data, a significant amount of research was then carried out in order to further define the reentry object requirements. Also, the recommended government furnished motors and other motors of the same class were researched in an effort to determine their engineering characteristics. At the Kickoff Meeting, which occurred at the conclusion to Phase 0, an interactive QFD was presented, which allowed the team to gain valuable insight into the customer requirements and target values. The QFD is a systems engineering tool that allows one to map all of the customer requirements to the engineering characteristics. This helps to identify the critical engineering characteristics in the design and the important trade studies 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 the most 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 Selection Phase 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 to the ALFT design problem. First order modeling and simulation tools were developed during 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 of each concept for the two representative missions. The results of these modeling efforts were then used to select a family of ALFT missiles from the feasible options. This was done using an interactive Multi-Attribute Decision Making (MADM) tool with industry input at a System Requirements Review that took place at the conclusion of Phase I.

Detailed Analysis The final phase of the design process, Phase II, focuses on the higher-fidelity analysis of the chosen concepts. Additionally, a key component of this phase was the design of the front section of the missile. The composition and design of the reentry object and avionics 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 missile construction were all examined. At the conclusion of the design, verification and validation will also be conducted on the ALFT concept to conform that it meets all of the requirements 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 and TRD 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 create low-cost target systems. These systems will be capable of completing missions of various trajectories and payloads. These systems must allow for short lead times using an all up round concept of operations, where the complete missile stack can be delivered fully assembled from the manufacturer. This “ship and shoot” method minimizes the assembly and 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 2 Range 1000 km 2500+ km

Number of Stages 1-2 2-3 Launch 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 Apogee Survival Altitude 100 km 40 km

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

Mass 400 kg 400 kg plus the mass of all reentry and range extension

components

TABLE 3 - AVIONICS SECTION SPECIFICATIONS

Parameter Delivery Order 1 Delivery Order 2

Associated Objects No Yes

Size N/A Four 8 cm diameter or two 12 cm diameter

Total Mass N/A 40 kg


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Flight termination and active guidance subsystems are driving requirements for cost. The ALFT system should use alternatives where possible that eliminate the need for these subsystems. 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 meet each delivery order. Some of these motors are from current United States government stockpiles. These motors are available at zero acquisition cost to the ALFT project and will 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 Oriole

Trident C4 3rd Stage Castor 1 Improved Orion Castor 4, 4a, 4b

ATACMS Orion 38 Patriot (PAC-3) Orion 50

SR19 Orion 50xl M57 Orion 50sg

Mk11 Mod5 Talos ASAT Stage 2 (Altair 3)

To best accomplish the requirements outlined, the ALFT shall consist of a family of vehicles with each missile satisfying specific missions or launch methods. This shall serve to ensure maximum system capability by not confining a single vehicle with the duty of fulfilling the broad array of missions outlined. Numerous constraints exist to limit and guide the ALFT design. Any solution to the requirements must be in accordance with the guidelines outlined by applicable documents such as customer supplied SOW and TRD. Also to be considered are applicable range safety documents, environment regulations, and international treaties. A detailed breakdown of all of the requirements as well as the team’s 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, it was necessary to eliminate combinations that were physically incompatible or that violated the requirements of the project. To do this, restrictions were placed on various aspects of the missile to ensure each option was viable. Then, an interactive tool called an Interactive Reconfigurable Matrix of Alternatives (IRMA) was created to help visualize 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. The first related to the launch vehicle. These restrictions ensured that the rocket would be able to physically launch the required payload and that each stage would be compatible with the others. First, any candidate missile needed a thrust to weight ratio of at least 1.2 for 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 consecutive stages to no less than 0.7 and no greater than 1.4. This constraint ensured that successive stages would fit together without creating excessive drag or necessitating a bulky interstage section. The second category of constraints related to the avionics section and reentry vehicle. It accounted for the requirements of each mission as stated in the provided documents. The avionics section can be separating or non-separating and contain 2 AOs or 4 AOs. The reentry object can be non-separating or separating, conic or bi-conic. The post-apogee survival altitude can either be 150 km, 40 km, or 0 km; this post-apogee survival altitude can 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 a combination of the two. This propulsion system can be powered by hypergolic propellants or a compressed gas Reaction Control System (RCS) system. Finally, the target range can be 1,000 km, 2,500 km, or maximum range on either a low or high trajectory. The filtering on this section of possibilities is directly linked to the requirements. The avionics section and reentry object do not need to separate for a 1,000 km mission; this mission’s avionics section does not carry any associated objects or a maneuvering system. The 1,000 km mission specifies a conic reentry object. The 2,500 and maximum range mission require a separating avionics section with associated objects and a separating reentry object. This reentry object must have a maneuvering system; the committee’s requirements specified a propulsive reentry system. This reentry object must be capable of a post-apogee survival altitude of 40 km, and must have a bi-conic shape. From these constraints, the IRMA tool was created, as shown in Figure 6. This tool allowed 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 is affected. Figure 7 shows the IRMA with some options chosen, and others eliminated by the compatibility 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 possible options was reasonable. Then, these options could be run through the simulation environment 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 large design space down to 1187 options. The feasibility and performance of these missiles must then be quantitatively evaluated. This evaluation is conducted using an extensive physics-based modeling and simulation suite of tools, comprised of seven different disciplinary tools linked together in the MATLAB environment. The seven disciplines featured are propulsion, geometry, aerodynamics, trajectory, range extension, thermal, and computer aided design (CAD). The flow of data through this modeling and simulation (M&S) environment is illustrated in Figure 8. Each of the seven missile disciplinary 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 given missile, which is calculated from the launch trajectory analysis and passed to the trade-off tool. The maximum range determines if the missile is able to meet the delivery order requirements 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 range extension requirements. If the range extension can in fact be met for a given missile, the weight of the necessary propellant and thermal protection system (TPS) are computed and also used as discriminators. The reentry trajectory code computes the burn time and in turn the amount propellant used, while the thermal analysis code determines the type and weight of the TPS. These metrics of range, propellant mass used, and TPS weight along with other launch vehicle and cost properties are then used to populate an interactive trade-off tool used in the downselection process.

FIGURE 8 - MODELING AND SIMULATION ARCHITECTURE


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Propulsion The 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 ten commercial boosters. This data was obtained from extensive research into public domain sources. Even after this research, a few details relating to some of the motors that are currently active in the military could not be found. This information was then interpolated from available data on similar systems in order to appropriately characterize these motors.

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

Nozzle Diameter Determination Nozzle exit areas are important in the calculating motor performance. However, they are also extremely difficult to obtain from public domain sources. As a result, it was necessary to calculate these areas from available pictures of the missiles. An example of this calculation is illustrated in Figure 9, which shows the motor and nozzle areas for the ATACMS missile. From this picture, the ratio of the nozzle diameter to the motor diameter 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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Motor Characteristics Summary Using the methods mentioned above, the necessary characteristics were compiled for all of the motors of interest [5] [6] [7] [8]. Dan Pickering of Teledyne Corporation also provided more accurate data for the Terrier, Oriole, and Improved Orion. A summary of the motor characteristics is found in Table 5.

TABLE 5 - MOTOR CHARACTERISTICS

Motor Name

Length [m]

Diameter [m]

Gross Mass [kg]

Propellant Mass [kg]

Thrust [kN]

Isp [s]

Burn Time

[s]

Exit Area Ratio

Terrier 3.93 0.46 966 680 279.3 247 6 0.85 Oriole 3.93 0.56 1174 976 92.5 289 30 0.81 Trident C4 3rd Stage

3.00 0.71 1863 1557 58.1 269 41 0.56

Improved Orion 2.78 0.36 395 293 25.3 228 22 0.49

ATACMS 2.00 0.61 1110 928 50.8 264 45 0.51 PAC-3 3.30 0.25 320 267 16.9 246 10 0.47 Talos 3.35 0.79 1990 1663 870.0 N/A 6 0.97 SR19 4.12 1.33 7032 6237 267.8 288 66 0.81 M57 2.70 1.00 1974 1657 84.0 273 64 0.49 Orion 50SG 7.60 1.27 13242 12154 485.1 285 73 1.21

Orion 50 2.65 1.27 3370 3025 118.2 292 73 0.42 Orion 38 2.08 0.97 985 782 34.6 293 65 0.72

Orion 50XL 3.58 1.27 4331 3915 153.6 290 73 0.72

Castor 1 5.92 0.79 3852 3317 286.1 247 27 1.00 Castor 4 9.07 1.02 10534 9265 407.4 261 54 0.61 Castor 4a 9.12 1.02 11743 10214 478.5 266 56 0.61 Castor 4b 9.20 1.02 11400 10000 430.8 281 64 0.61 Orbus 6 1.98 1.60 2749 2513 81.0 289 103 0.31 Altair 3 2.53 0.64 301 276 27.5 280 27 0.25

Geometry The geometry code determines total length and mass for a given configuration of payload and motors. It also generates the geometry and mass for the interstages and then creates a plot of points (distance from tip and radius) that the aerodynamic code Missile Data Compendium (MDATCOM) can then use. This code directly calls the motor database, which includes both the geometry and performance information for each motor listed in Table 5. In this way, the geometry code acts as the information center that other codes call to retrieve thrust and burn time information. At the beginning of the mission, the trajectory code calls the geometry


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module and initializes the missile with the correct reentry object and motors. At each of the subsequent stage separations, the trajectory code recalls the geometry module and creates 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 object research conducted by the missile design team. A summary of this research was provided to the MSTC shortly after the SOW and TRD documents were received. Included in this summary 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 version of the drawing in Figure 11 with a base diameter of 30” (0.762 meters). The two frustum angles and the tip radius would remain the same. An avionics section of 0.3 meters in length was appended to the bottom of the reentry object, completing the front section geometry. The bi-conic front section drawing is shown in Figure 12. This geometry is used 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), and top frustum angle (12.84°) were retained from the bi-conic design. The 0.3 meter avionics 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-conic reentry object). As noted in Table 1, the FS mass for DO1 is 400 kg, but the FS mass for DO2 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 components was created and is summarized in Table 6. More detail concerning the composition and component layout of the payloads will be discussed later.


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TABLE 6 - GEOMETRY MASS ESTIMATION

Component Mass [kg]

RO / AS Structure 400 Associated Objects 40

Propulsion Subsystems 15 Propellant 30

Thermal Protection System 10 RO Attitude Control 15

Miscellaneous Electronics 40 Total 550

Launch Vehicle and Interstage Assumptions All motors and payloads are assumed to have a constant density. Each of the interstages is assumed to have a constant length of 0.2 m and a constant density of 1500 kg/m3 [10]. The relatively short length of the interstages is used assuming that the nozzle is incorporated into the researched or provided motor lengths. Therefore, the interstage only needs to be long enough to provide room for the separating mechanism. Also, this interstage length worked consistently with the MDATCOM aerodynamic code while other interstage lengths generated persistent errors.

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

Missile DATCOM MDATCOM is a flexible aerodynamic code that enables quick evaluation of simple geometries. The shape of an axisymmetric body can be defined by a list of points (as described in the next section), which allows the missile geometry to be varied to match all 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 unknown reasons. If the failures were isolated, the points were interpolated from the surrounding values in the drag table. But when a given missile shape failed for all Mach and altitude combinations, the drag table was replaced by that of a similarly-sized missile’s table.


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FIGURE 14 - DRAG COEFFICIENT THROUGHOUT THE FLIGHT REGIME

Missile Shape Within 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 to supply a set of points defining the longitudinal (X) and radial (R) coordinates. The shape of the RO was constant, but the shape of the missile changed for each combination and for each stage of the flight. The X and R data for the missile shape was produced by the geometry code as previously explained. This data was then passed into the MATLAB wrapper 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 missile shape. The original data points are the green circles (only points of inflection were given). Then intermediate points were interpolated, starting at the center of the longest span. The interpolated points are shown as red X’s (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 missile to 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, azimuth and elevation, were provided as additional inputs. The propagation of the trajectory was then divided into three phases. The first phase, or the boost phase, simulated the launch of the rocket and integrated the trajectory until either the rocket reached 100 km altitude or the burnout of the final stage, whichever came later. The second phase, or midcourse and reentry phase then propagated the missile’s motion until it reached zero altitude. The final phase, or the range extension phase, simulated a burn performed beginning at apogee to attempt to reach the cross range and down range extensions required for the ALFT 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 shown using Google Earth in Figure 16.

FIGURE 16 - 1000 KM (RED), 2500 KM (YELLOW), AND MAX RANGE (GREEN) TRAJECTORIES

FOR A NOTIONAL MISSILE This module of the simulation was written in MATLAB like many of the other modeling tools. Various components were common to all three phases. The integrator used a fourth 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 oblate spheroid 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 is intended to be unguided, it was assumed that the rocket’s angle of attack would always be zero; therefore lift could be neglected. Also, jettisoning each stage after it burns out causes air turbulence which affects the rocket. However, this effect is relatively negligible and very difficult to model, and so was neglected. In this way, the only effect of jettisoning a spent stage is an instantaneous change in the vehicle’s mass. Since the rocket is initially rotating with the Earth, and the atmospheric forces are calculated relative to the planet’s surface, this phase was integrated in a topocentric frame centered at 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 at zero 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 aerodynamics modules respectively. These equations were propagated until the missile reached the edge of the atmosphere, defined as 100 km above mean sea level. However, if the rocket was still burning at this altitude, the equations were propagated until burnout and separation (if applicable) of the final stage. All position and velocity information was then converted into a non-rotating geocentric coordinate frame.

EQUATION 1 - BOOST PHASE EQUATIONS OF MOTION

( ) γμϕα sincos 2rmD

mTv T −−+=&

γsinvr &=

( ) ⎥⎦⎤

⎢⎣⎡ −+++= γμϕαγγ sincos1cos 2rm

LmT

vrv

T&

γθ cosrv

=&

spIgTm0

−=&

State Variables v: velocity magnitude r: radius γ: flight path angle (angle between velocity and horizontal) θ: range angle (angle between line from center of planet to rocket and similar line to launch site) m: mass

Other Variables T: thrust D: drag L: lift (assumed zero) µ: Earth gravitational parameter (3.986x105 km3/s2) α: angle of attack (assume zero) φT: thrust vector angle (assumed zero) g0: acceleration of gravity at Earth’s surface (9.81 m/s2) Isp: Impulse of motor


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Midcourse and Reentry Phases The midcourse and reentry segments of the missile’s flight were both modeled by the second phase of the trajectory module. These segments were combined into a single phase because the reentry vehicle possessed no aerodynamic maneuvering systems of its own 306H[11]. Due to this and the high velocity of reentry, the effect of the atmosphere on the 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 are actually negligible. The equations of motion for this phase, given in 307HEquation 2, are therefore based on Kepler’s laws of orbital mechanics. These equations were propagated until the missile reached zero altitude.

EQUATION 2 - MIDCOURSE EQUATIONS OF MOTION

rr

r v&&v3

μ−=

Range Extension Phase The requirements of the ALFT dictated that the missile must be able to obtain both a down range and cross range extension, so the third and final phase of the simulation modeled a deviation from the trajectory integrated in the first two phases using a maneuvering thruster on the reentry vehicle. The data for the thruster was provided by the propulsion module. By beginning the burn at apogee of the previously calculated trajectory, this phase of the simulation calculated the required burn time and mass of fuel to achieve the down range or cross range extension. In addition, it also calculated the requirements for achieving both simultaneously. The assumptions and equations of motion 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

rrm

Tr vv

&&v3

μ−=

An illustration of the range extension capabilities for a notional missile is shown in 309HFigure 17.


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

(GREEN) TRAJECTORIES FOR A NOTIONAL MISSILE Upon completion of all three phases of the simulation, the range and apogee of the missile 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 forces of 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, an optimizer was used with the trajectory module to find the appropriate launch elevations that would result in both the target range and the maximum range.

Code Evaluation To determine if the trajectory code performed as expected, it was evaluated using known trajectory data. This data came from the NASA Sounding Rocket Handbook that includes performance graphs for several sounding rockets [11]. The single stage Black Brant 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 an array of points varying the payload weight and launch angle. The physical characteristics of the Black Brant VC MK1 were added to the motor database, and then the trajectory code was iterated through all of the payload weight/launch angle points. The calculated trajectory data (impact range and apogee altitude) 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 known data. The worst case data point is at maximum payload mass (589.7kg) and lowest launch angle (76°). The results for this point deviate 5.4% for range and 10.1% for apogee.

Thermal Analysis One of the most difficult phases of ballistic flight is atmospheric reentry. An object reentering Earth’s atmosphere begins to experience drag and aerothermal effects below approximately 100 km, where the atmosphere begins to thicken to a substantial density at for an object traveling at high velocities. The temperature in the boundary layer around the reentering object can reach very high values, up to tens of thousands of degrees Fahrenheit, and because there is high surface shear, there is significant heat transfer to the RO 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 structural temperature lower then its failure point, and to keep the internal temperature lower then the payload’s 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 forms of TPS are passive and active systems. Active systems in general consist of some sort of insulating material that is integrated with a system that keeps that insulating material below a certain temperature. Such cooling systems include refrigeration, cryogenic fluids, and flash evaporation. Active systems 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. The downfall of active systems is that they are complex and heavy because of the need for storage of refrigerant fluids, as well as the movement of fluids through the insulating material. This adds cost, weight, and uncertainty of success. Passive systems consist of some sort of insulating or ablative material. Passive systems absorb and remove heat from the RO structure without the benefit of moving parts or additional systems, and so are desirable when possible because of weight and complexity savings. For this reason, passive systems, using either thermal soak or ablative technologies, were chosen for the TPS of the ALFT. Assuming the RO structure is strong enough to survive the thermal environment during launch, for the non-separating conic RO DO, aerothermal heating is not an important issue because it is only required to achieve a survival altitude of 150 km, and so no thermal analysis was done for the conic RO. For the separating bi-conic RO delivery order 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 analyze the reentry thermal environment and the type and amount of TPS required to allow the RO to survive to the desired survival altitude. The design tool selected for TPS design and sizing is NASA’s MINIVER, an aerothermal analysis and conceptual design tool. In addition, a zero-order conceptual analysis was written to supplement this analysis. Both analyses used a common TPS database from which to size and select.

Thermal Protection System Database The TPS database used for the TPS sizing was developed using NASA’s TPSX Material Properties Database, Web Edition Version 4, which was developed by the Thermal Protection Materials & Systems Branch at NASA Ames Research Center, along with NASA Langley Research Center, and the NASA Office of the Chief Engineer. It consists of an extensive list of TPS types and properties, drawing from the NASA Ames Thermal Protection Materials database and the NASA Johnson Space Center PathFinder Materials database. In order to be of use in the analysis, the density, specific heat, thermal conductivity, emissivity, and single use maximum temperature limit of each material needed to be available. From the TPSX database, 49 materials had the necessary information. Figure 19 shows a screenshot of the assembled database. It should be noted that this database also includes ablative materials. The single use maximum temperature given for these materials is actually the temperature at which the material begins to ablate. For the degree of fidelity required for the conceptual design of the ALFT, this use of the ablation temperature actually treats ablative materials as pure insulators. If the analysis selects an ablative material, it will then oversize the ablative material, but not grossly so. The result should 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 NASA TPSX MATERIAL PROPERTIES DATABASE

Zero-Order Conceptual Analysis The zero-order thermal analysis tool was written in MATLAB and was based heavily on the 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 the TPS material, the maximum temperature allowable at the RO structure surface, radius of the RO nose, velocity, and density at the corresponding altitude.

Key Assumptions Several key assumptions are used to simplify the analysis. The aerothermal effects experienced by the RO during launch are taken to be negligible compared to the heating that the RO is subjected to during reentry. Also, the nose is assumed to experience the highest temperature. To simplify the calculation of the weight of the TPS, it assumed that the TPS thickness is constant over the reentry surface of the RO, which is equal to the thickness of the TPS required to withstand the heat input at the RO nose. The TPS mass is assumed to absorb heat uniformly. The atmospheric density was calculated using the standard atmosphere model.

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

Higher Fidelity Analysis MINIVER was used in conjunction with the ASDL’s Thermal Management System Sizer (TMSS) for the higher fidelity analysis. It has been used in the past for both government and 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 a MATLAB code written and developed by Irian Ordaz of ASDL that generates the MINIVER input geometry and uses the MINIVER outputs to size and select TPS type using 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 generates streamline information from the VSP geometry. MINIVER uses this streamline information 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 best TPS 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 STREAMLINE GEOMETRY (B) TO MINIVER THERMAL ANALYSIS (C) AND TMSS TPS SIZING (D)

CAD For this study, Dassault Systemes’s Computer Aided Three Dimensional Interactive Application (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 spreadsheets and assign variables to any parameter within the model. This feature enabled rapid updating 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 select the 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 six M&S tools, it is possible that missiles with structural or aeroelastic problems could appear to perform well. By visualizing the options, these problems could quickly be eliminated. 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 from one 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 the missile FS. The purpose of this coarse model was to determine how to best organize the required subsystems inside of the RO and AS. Such items included the associated objects detailed in the TRD, a propulsive motor, a separating avionics section, sensory and telemetry gear, and a gyroscope for stability.


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Concept Selection After obtaining all of the results from the modeling and simulation environment, a significant amount of data was compiled. This data was then analyzed and put through a MADM 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 delivery orders- those cases that cannot meet the range requirements are discarded in a first pass filter). The judgment is parametric because of the influence of user-inputted weighting scenarios. The tool and its architecture are described below, followed by the final concepts 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 in which all of the important characteristics of the booster combinations are considered. Such an environment has been created in Microsoft Excel using a combination of worksheet 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 Solution In 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 initial form and normalizes and weights the data. Then, it picks out the positive and negative ideals and calculates the distances from these ideals for each alternative. Finally, the rank is based on which alternative is closest to the positive ideal and farthest from the negative ideal. The concept behind TOPSIS (shown only in 2-D here, but is actually multi-


31

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

FIGURE 23 - PARETO FRONTIER

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

Weighting Scenarios One 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, the tool will run anytime anything is changed on the sheet, thus creating an environment in which trade-offs can be discussed in real time. Essential to the form and function of TOPSIS are the weighting scenarios that are to be implemented 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 new and different set of rankings for each delivery order.

Delivery Order Commonality After 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 important characteristic of being “affordable,” it is necessary to consider the common boosters that are ranked highly in the TOPSIS results, to see if it is possible to ‘stack’ boosters or use the same boosters to perform all of the delivery order missions. This is done by calculating all of the possible common boosters in a given data-set using MATLAB, and then taking the rankings of each from their TOPSIS result, and adding them together, thus giving the best family of rocket boosters overall. This is a useful but not necessary step in picking out the final down-selection result.

Architecture of the Tool Figure 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 the spreadsheet and run through TOPSIS. Then, the results for the pre-determined families of rockets are assigned and re-sorted. This will yield a final concept list from which one can pick a family of rockets that both fits the delivery order criteria and is affordable.

FIGURE 25 - CONCEPT SELECTION FLOWCHART

Missile Downselection In the downselection process, several factors had to be weighted. Low cost requirements as well as transportability requirements were able to be weighed against the performance requirements. First the team only looked at the transportability and reliability by only looking at minimum stage vehicles. The Orion50sg and the Castor 4 series of motors can fulfill DO1 requirements as a single stage. Similarly, these motors will be a first stage for a


33

two-stage DO2 missile. However, these motors are commercial, and do not fulfill the need for a low cost target system. As a reference, Table 7 is shown below, showing the differing costs for each booster. The zero-cost boosters are Government Furnished Equipment (GFE) and the others are available commercially (COTS). Before actually performing any downselection, an “ideal solution” needs to be determined to compare each concept against. DO1 and DO2 will have the same ideal solution characteristics, but with some additional characteristics required to judge DO2 due to its atmospheric reentry requirements. These characteristics are shown in Table 8.

TABLE 7 - COST DATA Motor Name Cost to ALFT Program ($K) Terrier Mk12 0

Trident C4 3rd Stage 0 Improved Orion 0

ATACMS 0 Patriot (PAC-3) 0

SR19 0 M57 0

Mk11 Mod5 Talos 0 ASAT Stage 2 (Altair 3) 0

Oriole 600 Castor1 1,000 Orion 38 1,600 Orbus 6 2,000 Orion 50 2,400

Castor 4, 4a, 4b 3,500 Orion 50xl 4,000 Orion 50sg 6,000

TABLE 8 - TOPSIS IDEAL SOLUTION PREFERENCES

Factor Direction to Ideal Solution DO1 DO2 Number Of Stages Lower X X

Length Of Launch Vehicle Lower X X Weight Of Launch Vehicle Lower X X

Cost Lower X X Max Range Higher X X

RO Propellant Quantity Required Lower X RO TPS Thickness Required Lower X

Looking at only GFE motors by setting the cost dial to ten and everything else to zero (as shown in Figure 26), many GFE combinations emerged for both delivery orders. A Talos


34

first stage and either an SR19 or M57 second stage as well as an SR19 first stage and an M57 second stage can fulfill DO1 requirements. Other three-stage combinations of GFE work for DO1, but these were decided against because the TRD requires a two-stage missile for DO1. Similarly, SR19 and M57 based missiles are capable of meeting DO2 mission requirements.

FIGURE 26 - PROOF OF FUNCTIONALITY

Also, because a driving requirement is cost, it is imperative to look at the solutions with the best cost-factor, so in a similar fashion it is possible to look at the solutions with the least 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 are consistently at the top of the rankings: a Talos first stage, M57 second stage and a SR19 first 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 to discriminators in the operations and transportability of these two concepts. The SR19/M57 missile is a much more powerful concept with a predicted maximum range of over 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 elevation of 88.8°. A variation in range of 1,400 km in less than 3° of initial elevation means that the 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 a much lower sensitivity to deviations of initial launch elevation which makes it a superior choice given the requirement that this will need to be launched from the deck of a rolling ship. 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 for Maximum Range

[degrees]

Launch Angle Targeted to

1000km [degrees] Talos / M57 1294 64.0 74.6 SR19 / M57 2475 85.9 88.8

In addition to the favorable launch operations, the Talos/M57 is 5,174 kg less massive and 0.77 m shorter than the SR19/M57 making the Talos/M57 easier to transport, as shown in Table 10. The Talos/M57 concept’s superior non-performance characteristics make 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 4738 SR19 / M57 9.00 9912

When considering DO2, two GFE based concepts are consistently at the top of the TOPSIS results: Talos/SR19/M57 and Talos/SR19/SR19. Both missiles have approximately the same range so the main discriminators came down to operations and transportability. The Talos/SR19/M57 has a maximum range trajectory launch elevation of 75.2° and a 2,500 km targeted trajectory launch elevation of 83.3°. The Talos/SR19/M57 has a maximum range trajectory launch elevation of 84.9° and a targeted trajectory launch elevation of 87.7°. The Talos/SR19/M57 option is less sensitive 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]

Maximum Range [kg]

Launch Angle for Maximum

Range [degrees]

Launch Angle Targeted to 2500km

(High Trajectory) [degrees]

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

The Talos/SR19/M57 is also 5,000 kg less massive and 1.42 meters shorter than the Talos/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 12322 Talos / SR19 / SR19 14.32 17555

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

TABLE 13 - RO CHARACTERISTICS COMPARISON

Motor Configuration First Stage / Second Stage /

Third Stage]

Range Extension Motor Propellant Weight [kg] TPS Weight [kg]

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


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The Talos/SR19/M57 concept’s strengths in targeting and transportability as well as the M57 to avionics section interstage commonality make it the best concept for DO2. It is shown 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 have used 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 is the 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 design details for the missile can be determined. A more accurate estimate of the interstage weights 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 upon historical research of Titan and Minuteman missiles performed by the 2006-2007 Georgia Tech missile design team. For configurations where the interstage is connecting two components of similar diameter, this value is too large which results in a very short interstage. For these cases the half-angle is reduced until the minimum interstage length is 0.5 meters. Table 14 and Table 15 show the final geometry for each interstage for DO1 and DO2 respectively.

TABLE 14 - FINAL GEOMETRY FOR INTERSTAGES (DO1)

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

M57-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 0.55

SR-19 - Talos 16.7 0.9 Two stress calculations are performed to determine the acceptable wall thickness of the interstage frustums. The first calculation determines the compressive stress at maximum dynamic pressure (max Q) for each interstage. This calculation is based upon the drag force 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, max Q occurs at the same point for each flight profile. This point is six seconds into the flight at the end of the Talos burn. The second calculation takes into account the maximum allowable buckling stress due to axial 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 the selected material which is commonly used for solid rocket motor casings and interstages. For each interstage design the maximum allowable buckling stress is the driver in determining the minimum interstage thickness. Once this minimum thickness is determined, the frustum shell thickness is determined using the previously determined geometry and the density of 4130 steel. This value is then multiplied by 1.3 to account for the structure of the mating surfaces and the separation mechanisms. The new


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interstage mass estimates are found in Table 16 and Table 17 for DO1 and DO2 respectively.

TABLE 16 - INTERSTAGE MASS ESTIMATES (DO1)

Interstage Location Original Mass Estimate [kg]

Frustum Wall Thickness [mm]

Revised Mass Estimate [kg]

AS – M57 184 7 100 M57-Talos 190 7 102

TABLE 17 - INTERSTAGE MASS ESTIMATES (DO2)

Interstage Location Original Mass Estimate [kg]

Frustum Wall Thickness [mm]

Revised Mass Estimate [kg]

AS – M57 184 7 100 M57 – SR19 322 9 191

SR-19 - Talos 270 9 284 Figure 33 shows the drawing of the interstage linking the avionics section and the M57 (same for both DO1 and DO2).

FIGURE 33 - AS / M57 INTERSTAGE DRAWING


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Final System Design Overview With the final interstage dimensions determined, CAD models can be created for the DO1 and DO2 selections. These models are shown in Figure 34 and Figure 35, respectively.

FIGURE 34 - TALOS/M57 (DO1)

FIGURE 35 - TALOS/SR19/M57 (DO2)

The physical characteristics for each delivery order selection are shown in Table 18 and Table 19, respectively. The common Talos first stage, chosen for its excellent performance, yields very high thrust/weight ratios.


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TABLE 18 - TALOS/M57 (DO1) BASIC CHARACTERISTICS Parameter Value

Total Length 8.83 m Gross Liftoff Weight 4566 kg Maximum Diameter 1.0 m

1st Stage Thrust/Weight (sl) 18.3 2nd Stage Thrust/Weight (sl) 1.85

TABLE 19 - TALOS/SR19/M57 (DO2) BASIC CHARACTERISTICS

Parameter Value Total Length 14.47 m

Gross Liftoff Weight 12121 kg Maximum Diameter 1.33 m

1st Stage Thrust/Weight (sl) 6.91 2nd Stage Thrust/Weight (sl) 1.59 3rd Stage Thrust/Weight (sl) 1.75

Trajectory Performance The trajectory simulation can now be re-run with the revised interstage weights. The performance characteristics for each delivery order are shown in Table 20 and Table 21, respectively.

TABLE 20 - TALOS/M57 (DO1) TRAJECTORY PERFORMANCE

Parameter High Trajectory Low Trajectory Maximum Range (1488 km)

Launch Angle 76.5° 51.6° 63.3° Maximum Apogee 701 km 127 km 413 km

Maximum Dynamic Pressure 600 kPa 653 kPa 622 kPa

Total Flight Time 888 sec 403 sec 698 sec Time to Burnout

(2nd Stage) 71 sec 71 sec 71 sec

Free Flight Time 817 sec 332 sec 627 sec Burnout Altitude 104 km 45 km 78 km Burnout Velocity 1470 m/s 3040 m/s 3367 m/s

Downrange Distance at Burnout 41 km 82 km 72 km


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TABLE 21 - TALOS/SR19/M57 (DO2) TRAJECTORY PERFORMANCE

Parameter High Trajectory Low Trajectory Maximum Range (4166 km)

Launch Angle 83.7° 67.2° 74.4° Maximum Apogee 204 km 167 km 828 km

Maximum Dynamic Pressure 84 kPa 88 kPa 84 kPa

Total Flight Time 1803 sec 666 sec 1212 sec Time to Burnout

(2nd Stage) 138 sec 138 sec 138 sec

Free Flight Time 1665 sec 528 sec 1074 sec Burnout Altitude 259 km 64 km 161 km Burnout Velocity 5241 m/s 5288 m/s 5084 m/s

Downrange Distance at Burnout 110 km 254 km 223 km

Tables 13 and 14 show the different performance capabilities for each delivery order solution. The high trajectory would probably be the preferred method of launching each target, as the low trajectory launch angle is very shallow. The updated maximum range trajectory demonstrates the performance buffer that each solution offers. If a larger payload or longer range is required, each delivery order solution may still be capable of performing the mission. Figure 36 through Figure 39 graph the trajectory and mass change for each of the delivery orders.


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0 200 400 600 800 10000

100

200

300

400

500

600

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800

Distance (km)

Alti

tude

(km

)

FIGURE 36 - DO1 TRAJECTORY - 1000KM TARGETED (HIGH TRAJECTORY)

0 500 1000 1500 2000 2500 3000 3500 40000

100

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300

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(km

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FIGURE 37 - DO2 TRAJECTORY – MAXIMUM RANGE


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0 50 100 150 200 250 3000

500

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3000

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FIGURE 38 - DO1 MASS CHANGE (FIRST 300 SECONDS OF FLIGHT)

0 50 100 150 200 250 3000

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Veh

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FIGURE 39 - DO2 MASS CHANGE (FIRST 300 SECONDS OF FLIGHT)


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Thermal Protection System Design After completion of the trajectory performance modeling and simulation, conceptual design of the TPS could be conducted, using the MINIVER and TMSS modeling and simulation environment described previously. Because only the separating bi-conic is designed to enter the atmosphere, TPS analysis was only conducted on DO2 trajectories to determine TPS type and weight for each. The initial detailed analysis using MINIVER and TMSS was run on 806 reentry trajectories, using a sufficient fidelity of 300 grid points over the RO surface to achieve accurate TPS weights within ~0.5 kg. These TPS results were used as part of the down-selection to the final DO2 design. Once the final design and reentry trajectory were known, a high fidelity MINVER and TMSS analysis was run with 5600 grid points over the RO surface, yielding the results shown in Table 22. The thermal analysis showed the best TPS type to use as the Very Low Density Elastomeric (VLDE) Ablator, which is a silica-phenolic ablator found in NASA JSC’s PathFinder Materials database. Although the TPS analysis has given a specific TPS name, the fidelity of the overall M&S at this point in conceptual design indicates that the VLDE is probably only one of several silica-phenolic ablators that may be further explored in detailed design and testing in order to select the best TPS solution that takes into account not only mission success and weight, but also cost, availability, and experimental performance and reliability. The end result of this conceptual TPS design is to narrow the field of useable TPS forms substantially to the category of silica-phenolic ablator.

TABLE 22 - TALOS/SR19/M57 (DO2) TPS RESULTS

Parameter Minimum Range (2500 km)

Maximum Range (4166 km)

TPS Type Silica-phenolic ablator Silica-phenolic ablator Total Weight 8.5 kg 11.3 kg

Maximum Thickness 2.8 cm 3.6 kg Temperature

at Nose ~3590 K ~3590 K

Temperature Forward of Bend ~1922 – 2755 K ~2200 – 2755 K Temperature Aft of Bend ~1366 – 1922 K ~1644 K

Shown are graphical representations of the RO thermal environment determined by MINIVER for the minimum range of 2500 km (Figure 40) and for the maximum range (Figure 41) of 4166 km. Also shown are graphical representations of the TPS thickness distribution of the VLDE TPS as determined by TMSS for the minimum range of 2500 km (Figure 42) and for the maximum range (Figure 43) of 4166 km.


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FIGURE 40 - MAXIMUM SURFACE TEMPERATURES (OF) FELT AT EACH GRID POINT OF THE

RO DURING REENTRY FOR MINIMUM DO2 RANGE OF 2500 KM

FIGURE 41 - TPS (VLDE SILICA-PHENOLIC ABLATOR) THICKNESS REQUIRED FOR REENTRY

FOR MINIMUM DO2 RANGE OF 2500 KM


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FIGURE 42 - MAXIMUM SURFACE TEMPERATURES (OF) FELT AT EACH GRID POINT OF THE

RO DURING REENTRY FOR MAXIMUM DO2 RANGE ACHIEVABLE OF 4166 KM

FIGURE 43 - TPS (VLDE SILICA-PHENOLIC ABLATOR) THICKNESS REQUIRED FOR REENTRY

FOR MAXIMUM DO2 RANGE ACHIEVABLE OF 4166 KM


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Front Section Design The first order design of the reentry objects and avionics sections incorporates all component requirements from the TRD and SOW as well as any additional components necessary to perform propulsive range extensions in the case of the bi-conic reentry object. Table 23 shows all of the specifically listed requirements for components of the avionics section and reentry object for the conic RO. Duplicated requirements for the conic reentry object and avionics section (which function as one unit) are consolidated as avionics section requirements.

TABLE 23 - COMPONENT REQUIREMENTS LIST FROM THE TRD AND SOW FOR DO1

Reentry Object Avionics Section Hit Impact Location Measurement (HILM) System Thermocouples Strain Gauges

Comms. Equipment -C,S,X,Ku bands COMSEC Encoder Sensor Package -forward, rear, and aft -infrared and visible Navigation -GPS and IMU Command and Control System -fiber optic, RF, and hardwired Experiment/Instrumentation

Packages HILM System Thermocouples Strain Gauges

The bi-conic case requires all components listed under the conic case. It also requires additional components, which are shown in Table 24.

TABLE 24 - COMPONENT REQUIREMENTS LIST FROM THE TRD AND SOW FOR DO2

In addition to these components, several other component requirements were identified based on the bi-conic propulsive range extension requirement or other needs of the reentry object and avionics sections. These derived requirements are shown in Table 25 for the bi-conic RO. The only derived requirement for the conic RO was a battery in the avionics section.

Reentry Object Avionics Section Comms. Equipment -S, X, Ku bands COMSEC Encoder Navigation -GPS and IMU

Associated Objects Separation Control System


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TABLE 25 - ADDITIONAL DERIVED COMPONENT REQUIREMENTS CONSIST FOR DO2 Reentry Object Avionics Section

Battery Propulsion subsystem -rocket engine -propellant -pressurant -tanks, lines, valves Control moment gyro

Battery Pyrotechnic Separation Bolts

The bi-conic reentry object also requires components to enable propulsive cross range and downrange extension. A Leros 1B (as shown in Figure 44) bipropellant rocket engine provides 145 lbs of thrust at an Isp of 318 seconds. The Leros 1B is fully space qualified, having been used on spacecraft such as the Mars Global Surveyor, Mars Odyssey, and Mercury MESSENGER. The mixed oxides of nitrogen (MON) / monomethyl hydrazine (MMH) bipropellant used in the bi-conic reentry object is housed inside two equal volume titanium alloy tanks. The fuel system is pressure-fed by inert helium gas stored in a separate tank. A control moment gyro provides the attitude control necessary for exoatmospheric reorientation of the reentry object during cross range extensions.

FIGURE 44 - LEROS- 1B

Notional designs for the bi-conic reentry object and avionics section are shown in Figure 45 and Figure 46. The design for the conic front section is shown in Figure 47.

FIGURE 45 - BI-CONIC RO/AS BREAKOUT


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FIGURE 46 - BI-CONIC RO/AS ALTERNATIVE VIEW PARTIALLY EXPLODED

FIGURE 47 - CONIC RO/AS

All components were placed according to three basic principles. First, components must be placed in usable positions (e.g. the dispersing associated objects must not be blocked from being dispersed by other components). Secondly, geometric constraints must be met (i.e. the components must be arranged such that they fit within the designed reentry object and avionics section dimensions and volumes). Finally, weight should be centered axially for stability concerns.

The bi-conic case is designed as much as possible to be an extension of the simpler conic case. The avionics section for the bi-conic case requires a somewhat different layout than that for the conic case due to the inclusion of the associated objects. However, the bi-conic reentry object is designed as an extension of the conic reentry object. Since the geometric design of the bi-conic reentry object is based on adding a frustum to the cone of the conic reentry object, the additional components required for the bi-conic reentry object are designed to be incorporated within this frustum. A first order mass breakdown


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for the reentry object and avionics sections was previously shown in Table 6. A detailed diagram of the bi-conic front section components is shown in Figure 48. Table 26 provides a color coded key to aid in component identification.

FIGURE 48 - REENTRY OBJECT AND AVIONICS SECTION BREAKDOWN

TABLE 26 - FRONT SECTION COMPONENT KEY

Color Component Light Blue IR and Visible Sensor Package

Pink Experiment Section Red HILM Unit

Brown Battery Bright Green IMU Light Purple GPS Unit Dark Gray Antenna Set Light Gray Command and Control System

Yellow Associated Objects Blue Engine Black Pyros

Olive Green Propellant Tanks Dark Purple Pressurant Tanks

Burnt Orange Control Moment Gyro


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Launch Options The ALFT launch system complies with the launch specifications of nearly any ground launch base. The target vehicles missions will be conducted from or staged out of test ranges including, but not limited to, WSMR, RTS, Wake Island, Western Range, PMRF, WFF, VAFB, and EAFB. The ALFT will be capable of being transported and launched within 20 days of call-up from short or long-term storage condition. After delivery to the launch site, the system can remain on station for a minimum of 30 days in a standby condition. Once prepared at the launch site, the rocket will be ready for launch within 20 hours. ALFT was designed to withstand a variety of weather conditions maintaining at least 95% calculated launch target presentation availability. The systems are configured to withstand rainfall rates up to 10.2 cm/hr (4.0 in/hr) at wind speeds up to 64.4 km/hr (40 mph) when target is in its pre-launch configuration and a maximum snow load of 48.8 kg/m2 (10 lb/ft2) which is approximately 61 cm (24 in) of snow. The vehicle incorporates all environmental protection required for the system. In the event that a launch would be aborted the rocket is designed to return to ready for launch status within 6 hours of the abort. This includes replacement of expended items and recharging of batteries if necessary.

Shipping Logistics Three options for transport are feasible given the size and weight of both vehicles. The preferred method is shipment by flatbed trailer, given the cost efficiency and flexibility in reaching secluded launch sites. Also feasible is rail car shipping, though this method requires the existence or creation of infrastructure in the form of railroad tracks. Finally, air shipping via cargo aircraft is an option, although it is the most costly of the three methods and requires an airfield suitably close to the launch site. The missile can be shipped either in stages or assembled in the production facility and then shipped fully assembled. A missile shipped in pieces would be assembled and technically inspected at the launch site. However, this method is not preferred as it would require costly support equipment and trained personnel. Therefore the production method will be full assembly at the production facility unless otherwise specified. Special precautions will be taken in the shipment of the hazardous Hydrazine liquid propellant. As this substance is highly toxic, the system is designed to be fueled with the liquid propellant on-site. Another precaution for shipment is protection of the system against large shocks and vibrations through the use of specifically fitted packing materials and sufficient tie-downs. Sensors will be installed during shipment to ensure that the system is sufficiently protected. Packing materials and sensors will also monitor and protect the system against potentially harmful environmental conditions.


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Air Launch In the event that a ground launch is not a desirable option, air launch is also a possibility. Air launching has several benefits as compared to ground launching. It allows a vehicle to be launched from almost anywhere in the world and to achieve any desired inclination; it reduces the effects of drag by launching the vehicle from a thinner atmosphere; and it provides an initial velocity and altitude. The air launch methods that would potentially used are gravity air launch, depicted in Figure 49, and trapeze-lanyard air drop, as pictured in Figure 50 [21] [22] [23].

FIGURE 49 - GRAVITY AIR LAUNCH

FIGURE 50 - TRAPEZE-LANYARD AIR DROP WITH PARACHUTE STABILIZATION


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Conclusions An affordable and flexible family of target vehicles is needed to physically test the US BMD system. An SOW and a TRD requesting an ALFT system family were created in response to these needs. It was desired that the ALFT family use government furnished sounding rocket motors to fly both short and medium range missions. The 1000 km short range mission featured a conic reentry object, while the 2500 km medium range mission featured a maneuverable bi-conic. The ALFT family designed by the 2007-2008 Georgia Tech ASDL team used Talos, SR19, and M57 government furnished rocket motors. The short range concept was a two-stage missile using a Talos as the first stage and an M57 as the second stage. The medium range concept was a three-stage missile using a Talos as the first stage, an SR19 as the second stage, and an M57 as the third stage. The selection of these motors dramatically decreased the cost and schedule of the ALFT program. In addition to the design of the two launch vehicles, the Georgia Tech ASDL team also placed a large emphasis on the conceptual design of the two reentry objects and common avionics section. One large drawback to this study was that the team was limited to only public domain information. Much of the motor information was not publicly available and the information found was sometimes inconsistent or questionable. Information regarding reentry objects (especially bi-conic) was even more difficult to come by. Nevertheless, a large amount of research and engineering intuition was put into the study to ensure that the final results were reasonable. Although the chosen family is a logical selection, it is possible that a more optimal solution could have been found had the fidelity of available data been greater. The next step in the design of the ALFT system would be to obtain better data for the motors and more information regarding the reentry object attributes. Then, the same procedure that was detailed in this study could be equally applied to the new data. The team would be able to construct higher fidelity codes for the new data and rerun the physics-based analysis. The results of this new analysis could be examined with the ALFT downselection tool and compared to the initial results.


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Appendix A Requirements Verification Section Requirements Description / Response 2.0 Support Contractor shall identify the manpower, materials, services,

equipment, hardware, software, and necessary support equipment and other resources necessary.

Response: Qualitative Future detailed design would employ existing support and servicing

equipment where possible. For example, packing materials and tie-downs will be designed to integrate with current transportation methods, including flatbed trucks, cargo aircraft, and transport ships. The missile systems would be designed with current launch rail system interfaces in mind.

2.1.4 EMI / EMC The contractor shall develop and implement an EMI/EMC program

to ensure compatibility of booster system and payload systems from both intersystem and intrasystem design aspects. The contractor shall use MIL-STD-464A as guidance in implementing the program. The contractor shall use this program to ensure compliance of all EMI/EMC requirements during missile system integration, test range integration and mission activities.

Response: Qualitative Specification of equipment Electromagnetic Interface (EMI) and

Electromagnetic Compatibility (EMC) details is outside the scope of the preliminary conceptual design phase and will be determined during detailed design phase. Design of electronic systems and interfaces will meet interference control requirements as specified in MIL-STD-461 so that the overall system complies with all applicable requirements of MIL-STD-464A. This compliance will be verified by tests that are consistent with MIL-STD-461.

2.1.5 Range Integration External interfaces require coordination for utilization of range

integration and launch facilities. Response: Qualitative Missile external interfaces will be designed to utilize current launch

facility rail and support systems, with a specific focus on the launch facilities at VAFB, WFF, WSMR, RTS, PMRF, and EAFB.


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2.1.6 Mission Assurance The contractor shall develop a verification and validation (V&V)

Plan for the target missile system. Response: Qualitative Mission assurance will be provided by spiral verification and

validation (V&V) plan development. An initial V & V plan will maximize the use of computer simulations to minimize cost. A final V & V plan will be employed during the detailed design phase and initial production, with an attempt to minimize the need for both subscale and full scale testing. A launch preparation V & V plan will be developed in concert with the detailed design phase of the ALFT system to for customer use that will provide mission assurance for each individual launch.

2.2.1 Remote Launch Capability The contractor will also be required to provide a remote launch

capability. Response: Qualitative Detailed design of launch mechanisms is beyond the scope of

preliminary conceptual design but will be explored in detail during detailed design. The launch system will be designed to provide remote launch capability.

2.2.2 Missile System Tests The contractor shall define a test plan to ensure compliance with the

TRD. These tests may include engineering development tests, integration tests, qualification tests and acceptance tests.

Response: Qualitative As the missile systems are being developed, a continual test plan

will be employed in concert with the V&V plans to ensure compliance with the TRD. These tests will include but are not limited to engineering development tests, integration tests, qualification tests, and acceptance tests. Missile system testing will maximize the use of computational simulations to minimize test, but the use of physical subscale and full-scale system components will be required as design proceeds towards production and project completion. Testing will occur both on a component, subsystem, and full system scale to provide quality assurance.

2.2.2 Missile System Tests (cont.) The contractor shall provide any required test assets and personnel

to operate the test assets. Response: Qualitative The contractor will provide any systems, measurement devices,

computer programs, and materials required for missile systems testing, where feasible. Contractor personnel will also be provided


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to operate the test assets, as well as to train customer personal in system use. Training materials will also be provided.

2.2.2 Missile System Tests (cont.) The contractor shall recommend which of the Centers or other

contractor provided facilities provides the best value for ALFT. Response: Qualitative The preliminary conceptual design has been conducted with center

and facility flexibility in mind, especially with those indicated as important, including VAFB, WFF, WSMR, RTS, PMRF, and EAFB. During the detailed design phase, facilities necessary for production, transportation, and launch will be more thoroughly explored and final recommendations will be made at that time.

2.2.2 Missile System Tests (cont.) The contractor shall include in the Master Program Test Plan the

planning for integration, installation, checkout and test of the target missile components and system to be conducted at the selected Center/facilities.

Response: Qualitative After completion of the detailed design, the contractor will provide

and include in the Master Program Test Plan the planning for integration, installation, checkout, and test of the target missile components and system to be conducted at the recommended facilities.

2.2.3 Instrumentation and Telemetry The contractor shall design, install, and test instrumentation systems

for target missiles necessary to comply with the requirements of the TRD and DO.

The contractor will provide fully qualified special purpose airborne and/or ground based control, instrumentation, and telemetry systems.

To the maximum extent possible, these systems shall consist of readily available, COTS components, which will meet the environmental specifications for the rockets on which they will be installed and shall be compatible with range systems and GFE telemetry stations.

Response: Qualitative During the detailed design phase of the ALFT, instrumentation

systems necessary to comply with the requirements of the TRD and DO, including the HILM Device, necessary antennae, GPS, and RCS control systems, will be designed, installed, and tested before delivery to the customer. The contractor will also provide fully qualified special purpose airborne and/or ground based control, instrumentation, and telemetry systems to be determined in detail


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during detailed design, testing, and production that will interact with the ALFT instrumentation systems. Where possible, these systems will consist of readily available COTS components, which will meet the environmental specifications for the rockets on which they will be installed and shall be compatible with range systems and GFE telemetry stations, in order to minimize system development and deployment costs.

2.2.4 Experiment Support The contractor shall design and integrate Experiment Subsystems

(ES) onto ALFT motors necessary to comply with the requirements of the TRD and DO.

To the maximum extent possible, these systems shall consist of readily available, Off-The-Shelf (OTS) components, which will meet the environmental specifications for the rockets on which they will be installed.

Response: Qualitative Detailed design and integration of Experimental Subsystems (ES)

will utilizes past designs of sub-systems, using Off-The-Shelf (OTS) components that are currently used for similar or the same launch systems. The ES will meet the environmental specifications for the ALFT.

2.3.1 Pre-Mission Integration and Test Support The contractor shall identify test procedures and test support at the

test ranges to include vehicle buildup/integration and range integration.

The contractor shall identify instrumentation and telemetry recording at the launch facility for subsystem and system level integration testing to include initial checkouts, simulated countdowns, and troubleshooting.

Response: Qualitative During the detailed design phase, test procedures and test support at

the test ranges, including vehicle buildup and range integration will be developed, and instrumentation and telemetry recording necessary at the launch facility for subsystems and system level integration testing will be identified. This testing will include initial checkouts, simulated countdowns, and troubleshooting.

2.3.2 Range Support / Launch Services The contractor shall develop a plan to provide missile range support

services. Response: Qualitative A plan will be developed to provide missile range support services,

especially through detailed design, testing, and production.


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2.3.3 Range Test Support The contractor shall provide test equipment required in support of

equipment maintenance and specific missile range requirements dictated by the mission.

Response: Qualitative Test equipment that is required to support equipment maintenance

and specific missile range requirements dictated by the mission will be determined and designed during detailed design, testing, and production, and will be provided.

2.4.1 General Logistics Support The contractor shall develop an approach for storage, packing,

maintaining, transporting and tracking Contractor Furnished Property (CFP).

The contractor shall program and plan resources to deploy to the government furnished range integration facilities and range launch facilities. The contractor shall provide the personnel required for the deployment, operation, maintenance and retrograde of ALFT equipment and ordnance.

Response: Qualitative Details concerning storage, packing, maintaining, transporting, and

tracking Contractor Furnished Property (CFP) will be developed through detailed design, testing, and production. A program and plan will be developed concerning resources necessary to deploy to the government furnished range integration facilities and range launch facilities. Personnel will be required for deployment, operation, maintenance and retrograde of ALFT equipment and ordnance.

2.4.2 Transportation The contractor shall provide packaging and shipping of the target

missile system equipment and ordnance to the government designated range or staging area; and provide all necessary documentation to effect timely delivery of all equipment to include all necessary hazard classification data.

To reduce cost and time at the range the contractor shall consider shipping the ALFT as a unitary “all-up-round”.

Response: Addressed in design – qualitative note below During detailed design and testing, requirements necessary for

packaging and shipping of the target missile system equipment and ordinance to the government designated range or staging areas will be determined and will be provided upon delivery. All necessary documentation to effect timely delivery of all equipment will include all necessary hazard classification data. Preliminary design of the ALFT system was conducted with a unitary “all-up-round”


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method of shipping in mind, greatly reducing transportation cost and time at the range.

2.4.3 Missile Systems Maintenance The contractor shall develop a maintenance plan for the target

missile system component, missile related items, and ground support equipment.

Response: Qualitative A maintenance plan for the ALFT system components, missile

related items, and ground support equipment will be determined in detail during detailed design and testing of the ALFT.

2.4.4 Ground Support Equipment The contractor shall ensure that all GSE (GFP or contractor

provided) required for integration and ground processing, test checkout, integration and launch of the target missile system are available at all locations including the launch site, to include any GFP provided telemetry systems.

Response: Qualitative All Government Furnished Property (GFP) or contractor provided

equipment required for integration and ground processing, test checkout, integration, and launch of the target missile system will be made available at all necessary locations including the launch site, including any GFP provided telemetry systems.

2.5.1 System Safety Program The contractor shall implement an effective system safety program

for the design, development, assembly, integration and flight test of the target missile system. The contractor will use DoD 4145.26-M as a guide. The contractor shall comply with local explosive regulation for all operations and facilities involving explosive items.

Response: Qualitative Details concerning a system safety program will be determine

during detailed design and will provide a useable system for the design, development, assembly, integration, and flight test of the ALFT. DoD 4145.26-M will be used as a guide, complying with local explosive regulation for all operations and facilities involving explosive items. As per DoD 4145.26-M, potential for any mishap that could disrupt the ALFT program, delay production, damage or destroy GFP and contractor material and facilities, cause injury to any personnel, or endanger the general public will be minimized to the greatest extent possible. This will be taken into account in every phase of testing, production, transportation, and system use.

2.5.3 Range Safety The contractor shall perform a thorough safety assessment of the


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system design and its impact on safe conduct of tests. The assessment shall address ground and flight safety requirements.

Response: Qualitative A thorough safety assessment of the system design and its impact

on safe conduct of tests and missions will performed, addressing both ground and flight safety requirements. This assessment will begin during detailed design and will continue through testing, production, and deployment.

2.7.1 Environmental Compliance The contractor shall comply with all federal, state and local

environmental laws, regulations, and policies for all activities defined in this Statement of Work (SOW), whether conducted at government or contractor facilities.

Response: Qualitative All federal, state, and local environmental laws, regulations, and

policies for all required ALFT activities will be complied with, whether conducted at government or contractor facilities. An analysis will be conducted during detailed design to identify the environmental impact of production, testing, and deployment.

2.7.2 Operating / Non-Operating Environments The ALFT motors and its subsystems and components during

shipment, handling and pre-launch operations may encounter a variety of environmental conditions. The handling and pre-launch conditions are applicable to the ALFT booster, subsystem, or component without shipping container(s). ALFT design must allow for protection when exposed to the environmental conditions in the TRD. A summary of these environments is listed below:

1. Humidity: During shipping and handling, relative humidity from 0 to 100 percent with conditions such that condensation may take place in the form of water or frost. The contractor shall provide the capability to measure and record humidity while the ALFT is in transportation.

2. Temperature: Surrounding air temperature ranging from a minimum of -51°C (-60°F) during air transport or during ground handling to a maximum of +49°C (+120°F) during sheltered ground conditions and +71°C (+160°F) under unsheltered ground conditions. Areas protected from direct sunlight are defined as sheltered conditions. The contractor shall provide the capability to measure and record temperature while the ALFT is in transportation.

3. Vibration: When packaged or otherwise prepared for shipment, ALFT equipment shall withstand or be protected against transportation environments determined by analyses of handling and shipping conditions.


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4. Transportation Shock: Packaged equipment designs shall meet the requirements dictated by the contractor’s analysis of expected handling and shipping operations. The contractor shall provide a system to measure and record shocks encountered during transportation for each shipment of the ALFT.

5. Sand and Dust: During shipping, the ALFT shall withstand exposure to windblown sand and dust particles. During handling, the ALFT shall withstand dust concentrations of 6 x 10-9 grams/cm3 and particle sizes of 0.0001 to 0.01 mm diameter.

6. Rain: The ALFT shall be able to withstand exposure to rainfall rates up to 10.2 cm/hr (4.0 inches/hr) at wind speeds up to 64.4 km/hr (40 mph) during shipping and when target is in its pre-launch configuration.

7. Salt Fog: The liquid booster shall withstand exposure to a salt fog environment during shipping and handling.

8. Snow: The ground handling and transportation environments consist of a single storm snow load of 48.8 kg/m2 (10 lb/ft2) maximum (approximately 61 cm (24 inches).

Response: Qualitative Off the shelf equipment will be used whenever possible to monitor

the environmental conditions as the ALFT is being transported. Packing will be designed sufficient to protect the target from harmful effects of condensation, temperature, vibration, jarring, particulate materials, and precipitation. The specifics of these materials will be developed during detail design and manufacturing.

2.8.1 Product Assurance (PA) Program The contractor shall implement a PA Program consistent with the

objectives of ANSI/ASQC Q 90 quality series. The PA Program shall include planning that defines the contractor's PA objectives and milestones, identifies management and technical resources required to achieve the PA objectives, and describes in detail the methodology and schedule to be applied to PA tasks.

Response: Qualitative The Product Assurance program will be developed in accordance

with ANSI/ASQC Q 90. This plan will include system and component level testing to guarantee the capability of the system upon delivery. This testing will include inspection of existing GFE motors to determine that their function has not been compromised due to shipping or long term storage.

2.8.2 Quality Assurance Provisions In Documentation The contractor shall prepare, implement, and maintain quality

assurance documentation which provides procedures for examination and testing of all characteristics of the product being


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developed, fabricated, and used to conduct flight missions which affect safety, function, reliability, and life of the product.

Response: Qualitative The development of quality assurance documentation is not

appropriate for a preliminary level of design and will be addressed at length during the detail design phase. At the current stage of design, however, certain aspects of a quality assurance plan can be outlined. Because the motors used in developing the target are existing motors with proven designs that are already fabricated, existing procedures can be adopted during the manufacturing process. The areas in which a plan would need to be drafted are the connecting interstages between the motors, and the reentry object. For these aspects a detailed plan for examining and testing the characteristics will be drafted.

2.8.2 Automatic Self-Test Capability The system design shall include provisions for automatic self-test

capability. Control of the system during final count down shall be such that maximum flexibility exists in stopping, starting, holding, or recycling the count at any time prior to lift-off. The system shall always be capable of a controlled or safe shutdown in the event of any emergency or loss of external or internal power.

Response: Qualitative During detail design an automatic self-test feature will be adopted.

Because existing motors comprise the design, the systems already in use at launch facilities for self-testing, countdown, and countdown recycle will be employed as much as possible to ensure reliability and cost effectiveness.

2.8.2 Storage Requirements The ALFT shall be capable of being stored without liquid

propellants (if so designed) for a minimum of 10 years. The ALFT shall be capable of launching within 20 days (plus transport time (TBD) if mobile launched) of call-up from the long-term storage condition. After delivery to the launch site, the system shall be capable of remaining on station for a minimum of 30 days in a standby condition. Once prepared at the launch site, the ALFT shall be capable of launch within 20 hours.

Response: Analysis All components of the ALFT will meet the above storage

requirements. Hydrazine liquid propellant will need to be loaded at the launch site.

2.8.2 Booster Preparations Final booster preparations for launch shall not exceed 8 hours. The

system shall have a calculated launch target presentation availability


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of greater than 95% including the reliability of the ALFT and support equipment in a variety of weather conditions. The system shall support launch a minimum of 20 hours in any given 24-hour period. During preparations and launch, any environmental protection required by any element of the system shall be provided by the ALFT.

Response: Qualitative The procedures for launching the ALFT will be addressed at a more

detailed phase of design. As an overview, the procedures will include systems testing and environmental verification. The timing of these procedures will be designed to ensure a final booster preparation time of not more than 8 hours. The system will be designed to be robust enough to survive appropriate levels of environmental conditions leading up to launch.

2.8.2 Aborted Launch The system shall be capable of recycling from an aborted launch

countdown. In the event that the abort occurs the system design shall be such that it can be returned to a state where another final countdown can begin within 6 hours of the abort. This includes replacement of expended items and recharging of batteries if necessary.

Response: Qualitative Procedures for recycling from an aborted countdown will be

modeled from the current procedures for recycling a countdown in use at launch facilities. During detail design a means of recharging batteries and replacing expended components will be developed.

Safety The contractor shall comply with the requirements of EWR 127-1,

the EWR Manual. Response: Qualitative Eastern and Western Range (EWR) Requirements will be complied

with during all testing and launching procedures. Hardware The contractor shall employ best commercial practices in the

design, manufacture and testing of all safety related hardware. The ALFT shall be designed to provide for maximum safety, consistent with operational requirements, during every phase of the life cycle (especially loading, storage, transportation, range operations, launch, and demilitarization).

Response: Qualitative Best commercial practices (GAI/T-NSIAD-99-116, Commercial

Best Practices and the DOD Acquisition Process) have been employed during preliminary design through use of advanced


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design methods, especially the use of physical equations to model target behavior rather than historical-based data.

Hardware (cont.) The design shall stress the use of low toxicity propellants,

insensitive components, low hazard pressurization systems, generous safety factors, failsafe controls, and other robust but low complexity safeguards.

Response: Qualitative Most propellants in use are low toxicity; the exception is two

hypergolic fuels used in the RO which are highly toxic. The ameliorating factor is that these fuels are used widely throughout the industry, including on manned vehicles. For this reason, the highly toxic fuel is deemed safe enough for use.

The insensitivity of components, details about pressurization systems, and control failsafes are outside the scope of preliminary design and will be addressed during a detail design process.

Hardware (cont.) The design shall consider the safety requirements of each

anticipated transportation mode and incorporate materials and features to allow routine approval by the appropriate authorities.

Response: Qualitative The safety requirements of various transportation modes are outside

the scope of preliminary design and will be addressed during detail design.

Hardware (cont.) Design and operation of missile-pressurized systems shall be in

accordance with the requirements of MIL-STD-1522A, Standard General Requirements for Safe Design and Operation of Pressurized Missile and Space Systems. The respective range safety office must approve all safety-related hardware.

Response: Qualitative *Helium, fuel and oxidizer tanks in the RO

Detail design of specific pressurized systems is outside the scope of preliminary design and will be addressed at a later date. At that time, MIL-STD-1522A will be referenced as part of the design process to ensure compliance.

Support Equipment Requirements Support Equipment (SE) encompasses all hardware necessary to

transport, to launch and to support all activities and operations for processing the ALFT at the factory and field sites. Field sites


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include the integration, assembly, and test site(s), as well as the launch pad, air craft and sea craft.

SE performs both mechanical and electrical functions. The contractor shall maximize commonality between the SE used in the factory and that used in the field as well as on the launch pad.

Every effort should be made to minimize support activities at the launch site thru the institution of an “all-up-round” CONOPS.

Response: Qualitative The specific support equipment to be used falls outside the scope of

a preliminary design and will be addressed during the detail phase of design. Throughout this process, every effort will be made to enable use of existing support equipment in both the factory and launch site. Design of SE for the factory will be addressed during manufacturing planning.


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[18] “Defence Science and Technology Organisation (DSTO).” 01 Jun 2008. <http://www.dsto.defence.gov.au/attachments/20070612raaf8156021_0070.JPG>.

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[20] “Designation-Systems.Net.” 01 Jun 2008. <http://www.designation-systems.net/dusrm/app4/plv.html>.

[21] Katabeyoglu,A., T. Falconer, B. Cantwell, and J. Stevens, “Design of an Orbital Hybrid Rocket Vehicle Launched from Canberra Air Platform,” AIAA Paper 2005-4096, July 2005.

[22] Sarigul-Klijn, M.and N. Sarigul-Klijn, “A Study of Air Launch Methods for RLVs,” AIAA Paper 2001-4619, August 2001.

http://www.dsto.defence.gov.au/attachments/20070612raaf8156021_0070.JPG


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[23] Sarigul-Klijn, M.and N. Sarigul-Klijn, “Trade Studies for Air Launching a Small Launch Vehicle from a Cargo Aircraft,” AIAA Paper 2005-0621, January 2005.