Modular Liquid Propellant Launch Vehicle Design

Derek Yoshi Honkawa1 and Mahdi Yoozbashizadeh2

California State University Long Beach, Long Beach, CA, 90275, USA

Amatuer rocketry has access to a wide variety of similar and well understood components for the quick assembly of solid rockets from the scale of the smallest model rocket to that of sounding rockets, but this wide catalogue does not extend to liquid bipropellant rockets. Instead of designing a single rocket from start to finish to achieve a specific mission profile, a Modular Launch Vehicle Architecture (MLVA) was created to provide a great deal of flexibility to vehicle design and integration as well as a wider range of performance parameters for the ever-evolving Reusable Launch Vehicle (RLV) market. With many complex components in liquid rocketry, the MLVA approaches rocket design from the perspective of manufacturability and in performance by simplicity, rather than the traditional approach of optimized mass fractions. In doing so, rockets produced with the MLVA process are generally cheaper and easier to create than it’s custom-built counterparts while the vehicles and subsystems themselves are easier to test and scale using subscale vehicles. The MLVA design process was refined and applied to the design of two launch vehicles with two vastly different mission profiles, one for affordable moderate altitude flight and the other with the goal of breaking past the Karmen line of 100km.

I. Nomenclature Isp = Specific Impulse OD = Outer Diameter ID = Inner Diameter RLV = Reusable Launch Vehicle MLVA = Modular Launch Vehicle Architecture CF = Carbon Fiber (composite) FAR = Friends of Amatuer Rocketry LOX = Liquid OXygen WFNA = White Fuming Nitric Acid FFA = FurFuryl Alcohol CG = Center of Gravity Tg = Glass Transition Temperature IPA = Isopropyl Alcohol RCS = Reaction Control System TVC = Thrust Vector Control TWR = Thrust to Weight Ratio BLEVE = Boiling Liquid Expanding Vapor Explosion

1 Graduate Researcher, Mechanical and Aerospace Engineering 2 Assistant Professor, Mechanical and Aerospace Engineering

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II. Introduction Liquid bipropellant rocket systems consist of elements that make it typically far more complicated than it’s

solid and hybrid counterparts. While solid and hybrid rockets can be built from off-the shelf components into a variety of forms to suit its purpose, liquid rockets typically need to be custom built. This is often a matter of necessity since any number of factors can have a cascading effect on the many delicate vehicle subsystems. The particular flow characteristics required of liquid rocket engines are difficult to manage unless it is considered in the vehicle design. This becomes a major issue in launch vehicle development as each rocket becomes a case of reinventing the wheel; There are widely different launch vehicles with inconsistent design paths. At its core however, bipropellant rockets are mostly identical in that they all carry two segregated propellants that feed into an engine. All of these rockets have to carry something to define its purpose, and need to be sturdy to survive the trip, but there remains no “correct” way to create the most optimized vehicle. The drive to improve performance or cost or capability has led to launch vehicles approaching wildly different design doctrines. This behavior shouldn’t be discouraged, as the field of liquid rocketry is still full of unknowns and unexplored ideas that deserve discovery. The following document will attempt to deconstruct the vehicle design process and develop a common set of interchangeable and versatile modules that can be applied to a wide range of known and unknown propellants and design innovations. This effort will culminate into a comprehensive modular design standard for small scale bipropellant launch vehicles intended for research and educational settings.

The motivation for developing a modular design standard was brought about from the development experience of the Prospector 17c and Beach 1 bipropellant launch vehicles at Long Beach. The vehicles were both designed with full-composite skins, ostensibly to reduce overall vehicle weight. Despite superior performance on paper, the full composite skins have demonstrated severe drawbacks in development time due to the time it takes to manufacture the custom skins and the poor consistency and reliability of the resultant laminate, especially with the additions made for mating sections, like bulkheads and couplers. Bonding or connecting with composites was fairly ineffective since epoxy bonding post-cure is always weak, and bolts or shear connections are prone to shredding the composite plys. Since vehicle integration and architecture is entirely dependent on a completed skin, the vehicle development is stuck on a linear path. This meant that if the vehicle’s composite structures suffered setbacks or errors, which it invariably did, the development timeline is extended, and other subsystems have to wait for it to be redone. While nominally heavier for its strength than fiberglass or carbon fiber, aluminum can benefit from high consistency, availability, and isotropy which makes it a precise and reliable material. Aluminum can also benefit from fairly tough bolted connections and strong weldability. It’s precision translates to lower safety factors and simpler connections, which brings the efficiency of the two materials closer together. The modular system takes advantage of aluminum’s consistent characteristics to reduce launch vehicle cost and weight by removing variables and steps from the development process, allowing composites to contribute in different ways where it is far more suitable where it cannot become a potential roadblock for vehicle development and testing.

By maintaining consistent design elements, the development time can be reduced significantly while providing a greater range of test data to draw from that can be directly applied to any new configuration. With modular vehicle design, a subsystem can readily swap between different types and sizes while minimizing its influence on other systems. To demonstrate this, a series of modular parts have been designed for a common architecture and applied to the overall design and fabrication of two 8” diameter liquid bipropellant rockets with vastly different propulsion approaches, but identical maximum impulse. The first of these, already named the Mark 2 as it was the precursor to the current effort, will demonstrate a rocket with storable hypergolic propellants and a blowdown fed ablative engine, designed with an emphasis on low cost construction and high volume efficiency, ideal for education and for testing new modules. The second is a more complex vehicle using cryogenic liquid oxygen and alcohol using a regenerative engine and a regulated pressure system. These two vehicles define the range of potential rocket configurations offered, and demonstrate how to mix and match the modules to suit a customer’s needs. This is important for the modular system’s purpose in allowing greater flexibility and expandability as the user gets more familiar with liquid rocketry, which makes this an ideal tool for amateurs and educators, who may not find much value or enjoyment in building a strictly guided kit.

An individual new to liquid rocketry can start with a small and simple system before expanding and experimenting with it as they develop as rocket propulsion engineers. The modular system makes use of commonly obtainable materials for it’s construction, with the primary structural and propellant storage element being 6061-T6

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aluminum tubes available from most metal stock suppliers. By relying on ordinary stock as the primary element, the user can freely alter the lengths of the rocket sections, changing propellant volumes or giving more space for internal systems and modules. Repairs and upgrades would also be simplified since parts and modules would be interchangeable. The modular system allows rockets to test subsystems between vehicles so a cheaper vehicle, like a simple blowdown or even a solid motor, can be used to test subsystems in flight before being implemented on more complex and expensive configurations. This could include recovery, thrust vector controlled engines, cold gas thrusters, internal guidance systems, or vehicle staging for mating vehicles together. Any of these systems can be mechanically prototyped using 3D prints since they are not integral to the vehicle structure. 3D printed parts may not be the final module, but it can act as a stand-in while the actual module is developed outside of the vehicle.

III. Architectural Options The design choices and calculations leading up to the MLVA system may not be included with this

document. The vehicle engineer should be conscious of the limitations and structural analysis required for the launch vehicle. The MLVA is a set of guidelines for designing rockets with an emphasis on improved manufacturability and modularity, not on the efficiency or strength of the components demonstrated or used as examples.

Diameter A. 4” (0.065” walls) B. 8” (0.125” walls)

Structure A. Milled Tubular Frame B. Monocoque Frame C. Composite Skins

Bulkhead Connection A. Radial Bolts B. Welded

Propellant Tanks A. Tandem B. Concentric C. Common Bulkhead

Tank ends A. Flat B. Ellipsoidal

Engines A. Regenerative B. Ablative C. Adiabatic

Propellants A. Storable B. Cryogenic C. Autogenous

Max Tank Volumes A. 8” - 19.6 gallons B. 6” - 4.9 gallons

Pressure Feed A. Blowdown B. Regulated Pressure

Attitude Control A. RCS Modules B. Fin Guided C. TVC

Insulation A. Uninsulated B. Sabot Insulation

Main Valve Control A. Internal Piston B. External Piston C. Rotary Actuator

Payload & Modules A. Cubesat Bay B. Inline Module

Internal Access A. Access Panels B. Slide-on shell C. Rotatable skin

Fin Construction A. Composite fin can B. Welded aluminum C. Sectional Pieces

Table 3.1. List of Basic Options for Launch Vehicle Configurations

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IV. Industry Study The MLVA is a product of the Prospector and Mark 1 development trees. Much of the essentials of the

vehicle designs are based on standard methods and procedures originating from the Prospector family of launch vehicles. Most of what makes up small liquid reusable launch vehicle (RLV) design is further derived from a mix of amateur solid rocketry and a downscaling of full scale space launch vehicles. These two types of rocket aren’t very similar, so a rocket produced by these two development paths tend to be an interesting hybrid. Solid rockets are typically narrow, but a liquid rocket is more squat as the internal tank volume management has great influence on design and efficiency.

Few launch vehicles for orbital insertion use gas pressurization except for notably Vector Space rockets, while other companies like Rocket Lab uses turbopumps. The MLVA, perhaps as a result of it’s very distant shared ancestry with Vector Space, uses entirely gas pressurized propulsion. The simplicity of the system works well with the design theology of the architecture. Gas pressurization doesn’t perform well with larger volumes of propellant however, since it needs to carry more gas for each extra volume of propellant, while a turbopump works off of the propellants themselves or the battery power available.

The Atlas rocket relies on it’s propellant tank as the main structural element when it’s pressurized. It’s strength is as a result of the internal pressure acting upon the walls and preventing buckling since buckling modes would have to reduce internal volume. Many rockets have separate propellant tanks and structures, but since combining them can result in not only a reduction in the number of parts, but also an improvement of strength, these roles may as well be combined when possible.

Rockets for orbital insertion do so with multistage rockets, though most amateur and small scale rockets are single stage and very rarely two stage. Many rockets attempt to increase the performance of a vehicle in order to achieve as much as possible in one stage. If a rocket can be designed with many reliable stages like the Saturn V, the vehicle can achieve greater range, and the individual mass fractions are less important.

V. Mark 2 RLV The Mark 2 Reusable Launch Vehicle is the first developmental model of the MLVA. It is a product of the

Mark 1 development cycle after it was shelved for what became the MLVA, and is intended to be a low cost launch vehicle for sounding rocket flights or as a second stage sustainer. It lies on the lower end of the spectrum in terms of complexity and engine performance, but can compete against much larger and more complex vehicles due to its high efficiency and low part count. Equipped with an ablative engine and a blowdown propulsion system, the Mark 2’s hypergolic propellants will start and burn with very few points of failure. Its room temperature propellants, White Fuming Nitric Acid and Furfuryl Alcohol, are high density and easy to contain. The vehicle doesn’t have to worry about boiloff and can stay for some time, making it ideal for a second stage where loading is more difficult and secondary to the first stage. The seals for the tanks and the metals of the vehicle will not be affected or compromised by the temperature either. The propellants are stored in a single pressure vessel divided into two concentric tanks such that they share pressure sensors and valves, while a piston divides the ullage volumes. The high propellant density and concentric tanks result in a short and compact vehicle with a total impulse of 9200 lbf-sec.

The original purpose of the Mark 2 was to be a second stage vehicle, and while it’s role is now that of a normal single stage vehicle, it still has many of the design considerations for two stage vehicles in mind. It is perfectly suited for upper stage flight since the diameter matches that of another MLVA rocket. The high density also helps to shift the CG upwards in a stacked configuration and reducing the fin stability requirements of the lower stage.

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Figure 5.1. Mark 2 in flight as a second stage vehicle. (render)

VI. Mark Space RLV The Mark Space (Working Name) is the second generation vehicle designed using the MLVA. Unlike the

Mark 2, the Mark Space has a full aluminum body and all radial bolted connections. The engine is also a regenerative engine fed by a regulated pressure propulsion system working with cryogenic LOX and Isopropyl Alcohol (IPA). The first test phase has downscaled tanks for a maximum 9200 lbf-sec, but will be extended in later launches to carry enough fuel to cross 100km apogee. Unlike the Mark 2, the regulated pressure system and stacked propellant tanks require extensive plumbing. The Mark Space places it’s LOX tank below the fuel tank, and the fuel feed line is run along the outside of the rocket. This produces drag, but avoids the complexity of running fuel through cryogens which would need insulation to avoid gelling or freezing. The placement of LOX below IPA also means the external feed line does not need to be insulated itself, but since LOX is more dense than IPA, the CG is lower than it would be with swapped positions.

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Figure 6.1 Mark Space in flight with shortened tanks, fins, payload bay and cold gas thrusters. External plumbing is shown. (render)

VII. MLVA Common Diameters A defining parameter of a launch vehicle is it’s diameter. Rockets are commonly assumed to need to be

narrower for better performance, but this is not always true. Moreso with liquid rockets than anything else, length can cause issues if width is not increased with it. Whereas a solid or hybrid motor grain add stiffness and doesn’t slosh around, the liquid rocket has to contend with a hollow structure and a majority liquid mass. This makes the structural concern for liquid rockets more difficult to address. Another major factor is the structural efficiency of the propellant tanks and their shapes. A long noodle shaped tank will weigh far more than an equivalent spherical tank, but the aspect ratio of the tank is limited by the rocket’s diameter. A wider rocket may produce greater drag, though it will begin to weigh less and measure a shorter distance. A shorter rocket generally undergoes lower stresses and bending torque from flight and doesn't require as much additional stiffening as a longer rocket. Long structural elements are more likely to fail to flexural or torsional buckling, and if the rocket is not perfectly symmetric, as most liquid rockets aren’t, it can fail to flexural-torsional buckling.

Many small scale liquid rockets have fallen victim to the allure of a long and narrow rocket. A long history of solid sounding rocketry from university solid rocket teams and the Experimental Sounding Rocketry Association have pointed towards narrower being better for the reason of less drag. The reduction of drag is a powerful mover in design for groups new to liquid rocketry, but this benefit cannot be directly carried over to liquid design. Liquid and solid rockets have two completely different performance movers. Liquid rockets become far more capable with greater propulsive and mass efficiency, while solid rockets rely on improving aerodynamic efficiency. This is because solid motors cannot achieve much greater propulsive efficiency by switching solid fuel grains. The grains

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are room temperature and stored similarly with similar densities. The solid oxidizers are such poor propellants that a solid motor burns much cooler and slower than most liquid rocket engines, and produce large volumes of waste and impurities. This cap on propulsive efficiency is what allows solid rockets to almost exclusively use adiabatic nozzles like graphite. With long and narrow fuel grains, a solid rocket can burn fast and provide high thrust, which helps achieve high altitudes by piercing through the atmosphere with its low profile.

A liquid rocket cannot afford to be so long and narrow, since it would mean storing its propellant in a very inefficient straw-like shape, while the engine combustion chamber will have to be long and narrow to fit within the small profile. All the propulsion hardware has to fit within the space available. Solid motors also tend to be hollow core, such that the propellants will combust throughout the length of the rocket, making high thrust solid motors simple to achieve. A liquid propulsion system is more similar to an end burning solid motor. It can only burn so fast, but if it needs to produce more thrust it needs more width. This width is the surface area for solid fuel, and the plumbing diameters and sizes for the liquid rocket. This type of engine becomes less efficient if it has to burn up all it’s propellant quicker, so it is better to burn slower and over a far longer period of time. Most commercial liquid rockets will have a far lower Thrust to Weight Ratio (TWR) than it’s solid counterparts, with the famous Saturn V rocket sporting a TWR at ignition of less than 1, but they burn for up to several minutes. As a result, a considerable portion of the vehicle’s propulsion is performed in thinner atmosphere or vacuum, where the squat rocket does not face as much drag. For unguided rockets launched at the Friends of Amatuer Rocketry (FAR) launch site, a liquid rocket will almost invariably require the 60 ft Baxter T rail, while a solid rocket would launch off the 20 ft Newman rails. The reason is that the minimum velocity for fins to become effective takes longer to achieve with the liquid rockets than a typical solid.

To substantiate this line of thinking, a comparison study was performed by comparing apogee for different diameters of space-capable rockets, factoring in not only the changing shape of the tanks, but the fixed lengths, the fixed nose cone aspect ratio, additional module mass, and the increased thickness of wider tanks to account for greater hoop stress. The program used to create the different size rockets is an internally developed program. From table 7.1, the ideal rocket diameter was 7 to 9 inches. Notably, a 6 inch rocket was much longer, though it needed thinner walls for the tanks. This would be counterproductive to produce since it would also make the structure weaker. A 10 inch rocket and above needed wall thicknesses greater than ⅛” thick to handle the hoop stresses, which increased mass, while not making the propellant tanks any shorter. The increased mass from tank material and nosecone size required more propellants, and thus more tankage. This cycle continued, resulting in a levelling off of rocket performance at higher diameters where larger vehicles start to scale up evenly, but never with the same efficiency as 8 or 9 inches.

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Table 7.1. Key parameters for different launch vehicles capable of reaching space. Percentages are based on comparison with 8”, the standard diameter for MLVA.

Despite 9” being a well performing dimension, an OD of 8” was chosen over it for several reasons that justified the reduced theoretical performance. The primary reason is the availability of 8” OD tubing with thin walls. The tubing used for the 8” modular rockets is obtained from OnlineMetals.com, and has 1⁄8” thick walls and online posted prices. This diameter tubing is also compatible with 8” Ogive nosecones available from Madcow Rocketry. This removes the need to manufacture a custom fuselage or nosecone as would be required for 9” rockets.

A diameter of 7” also performs well, but it is not much different than the wider 8” and is less available. The reduced drag does not make a significant difference, so the 8” diameter would be preferable since it also contributes to larger payload space and makes it easier to work on the inside for wiring and assembling the propulsion system. A wider diameter also allows for larger and more powerful engines.

An 8” architecture has other unique design considerations involves. Most commercial desktop 3D printers, including the popular Prusa i3, can print parts up to this diameter, maximizing prototyping potential. 3D printing is ideal for manufacturing basic couplers, bulkheads, modules, bays, or even printing out prototype hardware like engines. The OD of 8” parts is also safely within the limits for most lathes and mills.

VIII. Vehicle Main Load Resisting System. That Which Holds All Together The standard material in the modular rocket system is 6061-T6 aluminum alloy, a high strength to weight

general purpose alloy that is affordable and effective for rockets. By making most of the parts out of this same metal, many parts can be made from bulk metal or from scraps and the temperature gradients are reduced to fewer locations. Unlike steel or the more expensive 7075 aluminum, the surface of the metal is soft and prone to scratching, but it also helps make the metal easy to work with and modify with hand tools.

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The benefits of metal body tubes is the precision in which they can be manufactured and the predictability of it’s bolted connections. The isotropy and consistency of the material is in stark contrast to that of composite materials, which makes aluminum easier and more forgiving to work with for a modular architecture. To achieve more efficient structures, the modular rocket architecture may include milled tubular frames. Instead of using a monocoque aluminum body tube, the tube may be modified by milling out sections of the tube to turn it into a frame. This frame may be designed with truss and stringer elements in mind according to the loading condition. A good design will lean towards using mostly straight stringers formed by milling several rectangular paths in equal sides of the tube. A pure truss pattern, one without any axial elements, is not ideal since the angled elements will be curved about a tube. The curvature weakens the element as it is already in a bending mode. This curvature is of little consequence for axial elements, especially when well braced with an outer skin. This frame should then be joined with a strong composite skin which provides axial and bending strength, as well as all the torsional resistance and stiffness sacrificed by the reduced frame.

The manufacturing methods for creating the milled tubular frame can also be used to create access hatches wherever necessary, and these hatches may be sealed by a similar outer skin. The Mark 2 uses a partially milled frame, with one access panel open to the outside, while the rest of the tube is preserved. The hatch will be concealed by a light section of CF skin that is attached to conceal the open hatch.

To allow for the application of modern advanced composites, the modular rocket design can also make use of the same universal 8” aluminum tube to make composite skins of any type. Since the 8” tube is the same as that used for the rocket, it can make use of the same ID and OD as the aluminum section. A skin laid up inside the mold tube can be fit into the ID of the modular rocket, making it an internal coupler, and the outside can be used to make any length of body tube. An external skin layup was tested using a method provided by John Newman for Carbon Fiber skins, and produced an 8” ID CF skin with 4 layers. The skin is able to slide over the outside of the Mark 2 propulsion section like an external coupler, and then radially bolted into an existing 12 bolt pattern for the propulsion section. After this fit, a new partial bulkhead may be bonded to the inside of the skin against the end of the aluminum tube to provide additional axial strength.

Several process improvements were made to existing composite layup methods as part of the development of the Mark 2. Chief among them was the Carbon Fiber layup method, which needed to produce strong and lightweight skins. Conventional vacuum bagging methods would tend to produce wrinkles as the compression bunched fibers together, weakening the structure significantly and creating points of failure. The process is also labor intensive and expensive despite the often poor quality and consistency of the end product. Heat shrink wrapping would still produce some wrinkling or waviness, but also provided fairly low compression, resulting in a weaker composite. The nature of the wrapping and heating process also result in less consistent quality. Newman’s method, however, takes advantage of the unique properties of Carbon Fiber, it having a low and sometimes negative coefficient of thermal expansion, to produce a cleaner skin without vacuums or heat shrink wrapping. The process begins normally, with a waxed mold, dry fibers, and resin. The layup is performed and let set until it has gelled. In the skin made during development, the method used is an interpretation of the original method, so many details may deviate. During the layup for example, the fibers were rolled up onto the mold from beneath like a rolling pin as seen in figure 8.1 in order to tightly fit the cloth around the tube as it is impregnated with resin. The fiber was all in a 0-90 orientation with one continuous sheet that spans 4 revolutions ultimately appearing as it does in figure 8.2. While the skin may be weaker in torsion, it’s not a mode of great concern for this particular skin.

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Figure 8.1 CF plys rolled beneath the aluminum mandrel. The plys are impregnated before being wound, and the weight and movement of the mandrel keeps the layers taut.

Figure 8.2. All layers are wound and the skin is left to gel.

As the resin gels it becomes gummy and incredibly viscous. The resin is essentially set in its location while still very pliable. Unlike its liquid state, the resin is semisolid, and while the skin is now one object, then resin has yet to solidify. While in this state, the skin was moved into an oven and ramped up to a low heat, allowing some moisture to evaporate and the aluminum mold to expand. The carbon fiber remains the same and is compressed together against the skin. This compression is essential for the strength of the skin, but had the epoxy been workable, the layers would just slip away. The heat also accelerates curing, which meant the skin was cured in just 7 hours since the beginning. The lubricant was not compatible with hotter temperatures and evaporated through the surface, creating greasy bubbles. This will be remedied by switching to PVA fluid and using a gentler ramp to allow moisture to evaporate more slowly. A layer of film release can also be added to the outside to provide a cleaner and smoother finish.

The skin removal process was also improved, or rather simplified, from the traditional method used at Long Beach. Whereas before a skin had to be laid up with an additional bulkhead and pushed out with a series of blocks against a long threaded rod, an attempt was made to directly bond spare fiber to one edge around the edge as of the tube, while the other end of the mold was secured to the end of the table. The newly laid up fiber had a thick metal washer strung through and attached to a ratchet strap secured to the other end of the table. After propping up the ratchet to pull from the center, the strap was ratchetted, gradually pulling on the entire skin as it gripped where it was bonded with epoxy. The skin came out easily like in figure 8.3, and in the future, this material may be added before the skin is put in the oven while it is still gelled, to provide a much better epoxy bond. The end result can be seen in figure 8.4, and the skin can now be cut to size to fulfill whatever role is required of it.

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Figure 8.3. (left) CF skin with removal material shown, after having been removed from the mandrel.

Figure 8.4. (right) CF skin prior to being cut and refined into the upper skin section for the Mark 2 RLV.

Composite skins have shown themselves to be inappropriate for the lowest section of the rocket where the resin may melt or burn in the presence of internal combustion, resulting in structural failure. Even after a burn, the residual flames and heat can set fire or decompose the epoxy binder and permanently damage the structure. Unlike metal counterparts, a composite skin cannot support metal radial bolts as a primary load path since it’ll tear up the layers with the thread, which means a bulkhead must be laid up inside the skin and can suffer from misalignment or poor adhesion. If the bulkhead is misaligned, it will have to be redone from the beginning, delaying the rest of the vehicle. Composite bulkheads are a major risk for engine attachment systems and makes integrating thrust vector control systems more complicated as new loads will have to be accounted for. Metal structures have the advantage of being readily used as both tanks and as a structural element, while a composite tank needs costly and time consuming efforts to achieve the same. The high degree of custom work needed to develop a composite system makes it easy for errors to be introduced at any point of the process, and modification is complicated by the reduced efficiency of bonding with cured epoxy. While composites still have a welcome and preferred place in upper sections of the vehicles and for secondary elements, like the nose cone, fins, and support skins, the direction of relatively small liquid launch vehicles would benefit from turning towards a balance between metal, and specifically aluminum, and composite elements.

Assuming a vehicle is integrated and aligned with a metal structure or skeleton, the composite elements would be able to be added or modified at any time without having to open up or modify the core propulsion elements. A rocket can still take advantage of the strength of composite structures for any section above the propulsion system when the engine thrust is mostly distributed below, and the skins now have to resist the drag loads from the nose-cone. A thin composite skin can also readily be used as an aerodynamic skin in conjunction with the aluminum structure.

IX. Bulkheads and Module Attachments. Holding That Which Holds Together The distinctive characteristic of the modular rocket design is arguably in the radial bolt pattern and the

connection between bulkhead and body tube. The same pattern and attachment method is used throughout the rocket, manifested as the typical 12 x ¼”-20 bolt arrangement. The pattern can be applied for both sealing and ordinary mechanical connection, and they can be freely placed along the entirety of the outer aluminum body with no work beyond drilling the pattern and deburring it.

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Figure 9.1. Two 12 bolt radial patterns with temporary screws attaching the body tube to internal pistons for the Mark 2.

How easy this can be achieved depends on how easy it is to make interchangeable 12 hole patterns. With an indexing head, making radial holes with a mill sets the standard, and is easy enough to achieve with a part like a bulkhead or a piston. This process becomes less practical with the body tubes and composite skins, where length and pattern placement become issues.

This was solved towards the end of production of the Mark 2 with a series of 2 tools, one designed for 8”OD, shown in figure 9.2 and the other for 8”ID in figure 9.3. The 8”OD tool is meant for the aluminum body tubes, while the 8”ID is meant for skins. The ID tool was already proven a complete success when the CF skin was made for the Mark 2. A preset distance was printed into the tool and allowed the pattern to be transferred onto the skin with a handheld drill.

figure 9.2. 8” OD tool for drilling 12 hole radial pattern. ¼”-20.

figure 9.3. 8” ID tool for drilling 12 hole radial pattern. ¼”-20.

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Figure 9.4. Drilling radial bolt pattern into CF skin using 8”ID tool.

Figure 9.5. CF skin installed with bolts into body tube and piston.

X. Propellant Tank Configuration.

A configuration available for flat faced tank ends is concentric, as used by the Mark 2. Unlike stacked tanks, both propellants drain at the same point, with one propellant stored in a donut around the other. This is similar in benefit and concept to the parallel tanks used on the Saturn 1b and the Lunar Module. Whereas a stacked configuration requires that one propellant must travel the length of the other propellant tank in addition to the travel to the engine, a concentric tank will have similar distances to travel, which increases pressure losses compared to the other propellant. While this is not that critical for rockets with turbopumps which pull their propellants individually, a gas pressurized rocket can usually only afford to pressurize both tanks to the same level, making balancing pressure more difficult. The concentric tank configuration also benefits from the removal of any need to run feed lines through other tanks. If the central tank is not completely sealed from the outer tank, but rather only at the bottom, a piston can be installed in the central tank and the two sections can share ullage. A shared ullage volume eliminates the need to manage each ullage volume separately. These benefits allow the Mark 2 to be so simple, with fewer parts and tubing than most other bipropellant launch vehicles.

The concentric tank can be set to a fixed total volume, but the ratios of propellants stored can be changed freely without modification. This allows the Mark 2 to also support different propellants and it’s inner tank can be widened or slimmed to more easily suit the architectural needs of the drainage ports.

The concentric tank has critical limitations. Unlike stacked tanks, the two propellants make liquid to liquid contact across the inner tank wall which permit greater heat transfer than if it were a liquid to gas contact. With cryogenic propellants, one propellant can freeze the other, and insulation would be much less effective. The piston concept would not be possible with cryogens or fuming propellants in the middle, since they would have to release gas that could not be managed otherwise. The cold temperatures would make dynamic seals of the inner piston less effective, as elastomers will become rigid and shrunken. While the concentric tank could still be possible with cryogens, it would not be nearly as simple as with room temperature propellants, and a stacked configuration would be a safer choice.

Beyond a simple stacked configuration, the modular rocket architecture fits well with the use of common bulkheads. Since every part is aligned based on the radial bolt pattern, the bulkheads are machined with very precise reference points. A straightforward execution would involve a separate body tube for each propellant with a shared bulkhead to seal off each tank individually with two piston seal sets. Alternatively, the piston seals that are often used for pressure tank bulkheads also allow a common bulkhead to be installed a considerable distance from the edge of a body tube, a feature that was confirmed with the installation of the bottom tank end for the Mark 2 from the top of the body tube. With the ability to install tank ends within body tubes, a sequence of 3 tank ends, the top, middle, and bottom, can be pre-assembled and pre-aligned using an alignment tool into the internal configuration for two separate tanks. The whole assembly can be installed by sliding or pressing all three into the body tube. This process would benefit from the inclusion of internal stringers or baffles to distribute the installation loads. Unlike the

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other arrangement, the body tube is seamless and only one set of piston seals are needed, which reduces total part mass.

The quality of a propellant tank is important to mitigate the risk of BLEVE, or Boiling Liquid Expanding Vapor Explosion. BLEVE is a catastrophic and destructive event that does not need an engine to initiate, but can occur with any pressurized cryogen or nitrous oxide. The mechanics of a BLEVE are the simultaneous state change and expansion of a pressurized liquid that is made possible by the nature of phase states. Any pressurized liquid boils into a gas at different temperatures as pressure increases. Nitrous oxide, popular with hybrid motors, is liquid at almost 900 psi at room temperature, but a gas at atmospheric pressure. As one of these BLEVE liquids warm up, they will want to dwell at higher pressures, so LOX that is not fully insulated will release gaseous oxygen until it reaches a stable pressure for the given temperature. Many cryogens, like LOX, do not have a stable liquid pressure at room temperature. Should the vessels storing these propellants should fail, such as if a weld cracks at low temperature, the store liquid will not only want to rapidly boil into the atmosphere, but the high pressure gas will want to expand, creating a very dense gas explosion at a much higher pressure than the liquid was originally stored in. In one example, a vessel of refrigerated nitrous oxide at 300 psi suddenly bursts, and the resultant gas will not expand out as 300 psi into 14.7 psi, but rather as 900 psi. This turns an ordinary leak into a violent explosion capable of shredding the rocket apart and into various shrapnel. Such a BLEVE detonated an SDSU rocket on the launch rail, sending shrapnel a fair distance and destroying the Baxter T rail. This incident proved a fair reminder of the risks of cryogens and the steps that should be taken to ensure the propellant tank is sound. The failure of the SDSU rocket was believed to have stemmed from a longitudinal aluminum weld. To help prevent such a failure from occurring again, the MLVA has no provisions for longitudinally welded tubes, and the 8” body tubes used for the propellant tanks are extruded.

XI. Propellant Tank Ends For the ends of the tanks, a typical shape is ellipsoidal, which is ideally suited for reducing material usage,

while also providing fair drainage characteristics. The ellipsoidal shape is also ideal architecturally. When cut out of a cylindrical piston, where the middle is thinner and the walls tall, there is plenty of space for o-rings or radial bolts without having to support that waste mass in the middle.

Figure 11.1. Example of Ellipsoidal Tank End with weld-on attachment. All structural shells are attached with radial bolts as shown here. This doubles as an auto-centering feature which helps keep the vehicle consistently aligned to a reasonable extent.

For simpler fabrication or special cases, flat faced tank walls may be used. This is much cheaper and easier to produce than the ellipsoidal, since while ellipsoidal tank ends may be more efficient, they require more metal stock at the beginning. To meet the same strength of the ⅛” thick walls, only a 1⁄2” aluminum plate is needed to cap off the ends. Though the tank end may still bend, it will hold. These flat face tank ends are especially effective for the tops of tanks, where shape isn’t important for flow, and allows for easy modification to include tank support hardware and sensors directly to the tank, as with the Mark 2.

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Figure 11.2. Flat Tank End Concentric with piston type attachment. Differences in propellant volumes can be controlled by filling the chambers to different heights, providing a degree of flexibility. Not much variation is offered as this type of tank cannot be used with cryogenic propellants without modification.

The connections of tank ends is important for preventing BLEVE. Welded tanks are an attractive option since it produces the most efficient connection and seal for a propellant tank. They do not require special attention or care like an o-ring seal, but they are still unpredictable. Welding aluminum will especially weaken the alloy as it loses its temper, and the material itself is often described as temperamental and difficult to weld well compared to steel. The result is a joint that can appear sound on the outside, but possess a labyrinth of microcracks and weak points that can make a sudden appearance at high pressures or low temperatures, or even just after a successful landing. While welded connections are still viable, they must be approached with considerable caution, lest they prove a failure point for a catastrophic BLEVE.

XII. Engine Propulsion Options With the modular rocket design, a great emphasis is placed on common components that can be used for a

wide range of purposes and design parameters. This poses an issue with the user of the system who would most likely want to use a specific or a custom engine. The Mark 2 demonstrates the radial bulkhead and flange connection thrust plate. This configuration is ideal for conventional flange mounted engines up to 7.75” in diameter. The flat flange and the simple disk bulkhead is a generous platform to work with, and provides a place to place ports or mount TVC hardware. For the less adventurous, an LR101 can be comfortable integrated to this plate. Many shapes and sizes of engine can be tested with this attachment method, and this process is made even simpler with the standard ablative engine flange of the Mark 2. While combustion chamber volumes are very specific for each engine, the diameter can be fairly easily controlled to best suit the launch vehicle needs. The standard flange is flange mounted to the thrust plate and supports a radially sealed injector of set dimensions. The flange itself holds an ablative chamber of 3.75” OD. The chamber itself has to have an inner diameter of 3”, which provides ample depth for hotter or longer burning engines. The difference in wall thickness can be overcome with fiberglass/epoxy filler and turned to size.

As part of the research towards simplifying and modernizing engine fabrication, the first Mark 2 engine was built using 3D printed sacrificial mandrels and tools like that of figure 12.1. Instead of machining a two part aluminum mandrel, the 3D printed plastic would perfectly replicate the curvature which reduces mandrel mass and cost. This type of mandrel would not be recommended for heat cured phenolic resins, but cured perfectly with room temperature catalyzed phenol resorcinol resin. After the engine was cured, the plastic mandrel is removed from the inside in a labor intensive process with pliers and pvc pipe. Besides the low cost, the single use 3D printed method does not provide any particular benefits over the reusable aluminum, but it may still be appropriate for prototyping or short runs of engines.

Another experimental change was the fiber layers of the composite chamber, which was changed from a typical tape wound construction to a pseudo-filament-wound chamber. The pure silica fibers of the engine chamber were pre-formed into a woven tube which was flexible enough to seamlessly conform to the curvature of the engine,

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even at the throat, which can be seen to work in figure 12.2. The elimination of separate tapes should also help to reduce the flaking away of chunks of engine during firing.

Figure 12.1. Plastic Engine Mandrel, after having been spray painted, waxed, and lubed.

Figure 12.2. Silica Tube Sleeve Contour Conforming.

Figure 12.3 Cutaway and Top View of Ablative Engine Chamber Post-Cure Different rocket engine types can be installed in MLVA rockets, including regenerative and adiabatic

cooled engines. The Mark Space has a regenerative engine based on the MLVA architecture that has consistent diameters and adjustable lengths to suit the engine combustion needs. It also makes use of the body tube itself to join the sections of the engine together, which helps reduce the number of parts in the load path and makes the most of existing structural elements. Besides regenerative engines, adiabatic thrust chambers can be used for warm combustion for a simple low performance engine. Combined with the simple construction of MLVA rockets, a liquid rocket can approach the simplicity of the solid rockets.

As the MLVA uses body tubes, a solid rocket motor can be readily swapped into a rocket using the same attachment points and hardware using centering rings.

XIII. Vehicle Propellants With the goal of creating a more affordable high performance rocket, the use of storable propellants, those

which can be stored at room temperature, can eliminate many of the design challenges introduced with cryogenic handling and design. Despite being constrained to certain criteria, the benefits of a unified and modular rocket design is its flexibility. A modular rocket can be configured to a generic storable propellant combination such as Hydrogen Peroxide/Alcohol, or White Fuming Nitric Acid/Furfuryl Alcohol (WFNA-FFA) as used in the Mark 2.

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White Fuming Nitric Acid and Furfuryl alcohol are hypergolic, meaning they have a negative activation energy, and will combust spontaneously upon contact. Just like a monopropellant Hydrazine engine, an engine with these propellants does not need an igniter to begin combustion. This comes as a double edged sword, since hypergolic combustion is much faster than ignited combustion. While the spontaneous combustion prevents the chance of a hard start from propellants that haven’t heated up fast enough, it can also produce a hard start from propellants entering the combustion chamber too quickly before combustion can reduce flow rate through orifices, meaning the main valve must be calibrated to open slowly. From a cost perspective, these hypergolic propellants remove the need for igniters, while from a system perspective, the lack of an igniter removes a significant failure point in a launch or firing, and one that doesn’t need to be reset by hand if it fails. The reliability of a hypergolic system is its greatest strong point, which made it the system of choice for the engines on the Lunar Module. This same reliability makes the Mark 2’s hypergolic propellants ideal for upper stage vehicles.

There’s not much to say about Furfuryl Alcohol from first impression. It’s a niche byproduct of waste biomass, unlike the more popular hydrocarbons, and it’s only notable application is for use with WFNA in rocketry, which itself is very limited. It is remarkable that it has not found more extensive use however, since it has the highest density (at 1130 kg/m3) of any conventional rocket fuel, putting it on par with some oxidizers. Combined with WFNA and it’s high density of 1510 kg/m3, these propellants ends up being one of the densest bipropellant combinations in conventional liquid rocketry. This in turn allows a rocket fueled by WFNA-FFA to be very compact, as is reflected by the size of the Mark 2. While the specific impulse of WFNA-FFA is on the low end of the spectrum, the remarkably high density of the propellants allows far more propellant to be carried in the same vehicle, which is reflected in density-specific impulse.

White Fuming Nitric Acid is, like most oxidizers, dangerous and temperamental. WFNA is a high concentration (>90%) of Nitric Acid, beyond that of Red Fuming. It will react to aluminum, but this can be easily resolved by passivization of the tanks with a small amount of nitric acid beforehand. It’s also high corrosive, so it must be handled carefully, but is more stable than High Test Peroxide or even Nitrous Oxide.

XIV. Propellant Tank Volumes and Mixture Ratios Propellant tanks in MLVA are fully scalable depending on the desired total impulse or mixture ratios. The

locations of the piston ends and the distance between them are determined by the locations of radial bolt patterns on the body tube. This is in direct contrast to the typical welded unibody cylinder. A unibody tank can be mounted within the body tube and secured with bulkheads and centering rings, but with the aluminum body tube able to support over 900 psi of internal pressure, it is not necessary to have to make a separate pressure vessel.

The scalability of the propellant tanks are not just a feature for ease of manufacture, but also for testing and prototyping. A vehicle can be tested and launched with shortened propellant tanks, all other features being the same. Once the rest of the vehicle is tested, such as any guidance and avionics, or the rest of the propulsion system and engine, or the aerodynamics or recovery systems, the propellant tanks can be swapped out for a full scale launch with minimal further modifications. The primary differences in the vehicle flight characteristics during this transition is the aerodynamics, Thrust to Weight Ratio (TWR), and of course the apogee. Because the tank ends and the tank diameters are the same, the propellant storage and drainage characteristics has little reason to change.

XV. Propulsion Gas Pressurization Systems In addition to the core principles of the modular rocket architecture, the propulsion plumbing system relies

on using double compression fittings and O-Ring Boss connections wherever possible. It’s performance at high pressures and its resistance to loosening is far superior to AN and NPT fittings. NPT fittings are difficult to avoid however, and are often found standard with typical Pressure Transducers, relief valves, and Solenoid Valves. AN fittings are typical for external connections and have been found to still be acceptable, given it’s robustness. The downsides of NPT connections are best mitigated by reducing the number of connections to manage.

While gas turbopumps have no current place in the MLVA, gas pressurized feed can be simple and complex depending on the required performance. The simplest gas pressurization is a blowdown system like that found on the Mark 2. The blowdown pressurization has a fixed internal ullage volume. Ullage refers to the gaseous volume in a propellant tank. At engine start, the stored ullage pressurizes the liquid propellant as it feeds into the

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engine. The engine combustion drains the propellant tank, reducing incompressible liquid volume and increasing compressible ullage volume, causing the overall internal pressure to drop since the same amount of gas must now fill a larger and larger volume. While a blowdown system is simple and doesn’t require a separate gas tank and the plumbing to supply it, it also causes the engine to lose efficiency over time as the propellants feed at a lower pressure. Since the fixed internal ullage is necessary to feed the propellants, a larger proportion of the tanks internal volume must be kept empty, which can as much as double the length of the propellant tanks.

For longer burn times and when engine performance must be steady, a regulated pressure system like that used with the Mark Space is available. Regulated pressure is heavier and more complex, but improves performance in several ways. Constant pressure improves engine performance by maintaining it’s full combustion efficiency throughout the burn. Regulated gas is drawn from a high pressure source, which is more efficiently stored than with blowdown systems, and the ullage volumes of the actual tanks can be kept low. This means that less space on the rocket is used on empty gas, and this benefit stands out for larger propellant volumes.

The Mark 2 was tested for pressure decay at room temperature lab conditions with losses shown in Figure 15.1. This system, which uses viton™ o-rings in pairs at each end, had some difficulty integrating with the body tube, since while the pistons were perfectly round, the tubes were used as-purchased. The tubes as they were delivered by Online Metals were slightly out-of-round, and the extent of which varied with the batch and even across the tube. The pressure decay tests are important for a piston-based propellant tank, since unlike a welded tank, a slight pressure loss over time is nominal, and the effectiveness of a tank is determined by how small that pressure loss is at significant pressures. The in-lab decay tests are limited to low pressures for safety reasons, and gaseous nitrogen is used to fill the tank to a specific pressure, before the regulator is shut off. The results of the tests show an acceptable pressure loss of nearly zero for low pressure ranges. This test is useful for only major leaks and system compromise, where high pressure leaks would be dangerous. Further testing will be performed after this at the FAR site with LOX and high pressure nitrogen or helium to test the system under stress. The effectiveness of the Viton™ o-rings will perform different when in prolonged cryogenic conditions and high pressure, where the fluoroelastomer will harden. At the same time however, a well polished inner and outer metal surface will create a metal seal for as long as the outer shell cools faster than the inner piston. This can be a deliberate design feature if the outer shell is insulated while the tank ends are uninsulated or even heated. A steel piston would also shrink less than the outer aluminum, making it ideal for a cryogenic system. The lack of a temperature gradient between dome and steel fittings also reduces the potential for leaks at this contact, though this comes at a significant mass increase.

Figure 15.1. Pressure Decay test of piston sealed Mark 2 tank.

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XVI. Guidance and Stability Vehicle guidance and stability is a crucial consideration for a rocket’s mission profile. A rocket with just

fin guidance will quickly turn and spend energy on the wrong direction and make any recovery into a considerable effort. For most applications, fin stability is preferred for its reliability and simplicity. The MLVA has provisions for active guidance and stability options such as reaction control thrusters or thrust vector control. The swappable architecture allows an independent Reaction Control System (RCS) module to be placed as an addon. Unlike integrated RCS systems, the RCS module allows a vehicle to fly dumb, without the module, and the RCS can be added later.

Thrust Vector Control (TVC) is the preferred guidance method for most commercial boosters, and is an incredibly fast and powerful control method. TVC has the benefit of requiring no more fuel than is already available for main propulsion, and a large vehicle for spaceflight will require long burn times. The increased duration would result in increasingly prohibitive amounts of RCS fuel or gas, so the use of TVC or RCS should be carefully considered. TVC and any active guidance system requires considerable testing and precise hardware to perform adequately, and a simple MLVA vehicle provides an ideal analogue for testing TVC for a more expensive MLVA vehicle. TVC actuators should be integrated with the modular thrust plate, keeping all attachment points and hardware in one place and making it easy to swap between a guided and unguided engine, as well as removing the engine without disassembling the subsystem.

XVII. Propellant and Vehicle Insulation Insulating a vehicle for cryogens can add a not inconsiderable dry mass to a vehicle. For the modular

architecture that relies on the body tube as the propellant tank, there is no space for insulation to be installed around the propellant tank without going on the outside. For a launch vehicle with aluminum walls, heat transfer would make propellant boiloff a serious issue. High insulation can be provided with sabot type insulation, which can provide the most insulation while adding the least mass to the vehicle. Unlike many launch vehicles which either carry their own insulation or continually replenish it’s cryogens to replace boiloff, a sabot insulation is shed at takeoff and does not contribute to launch weight. Such a system is planned for the LOX-IPA Mark Space rocket.

Sabot insulation is a piece of ground support hardware that has two key features: It hugs the outside of a tank in two or more form fitting pieces, and it is not held on to the rocket with any great security. The sabot insulation planned for the Mark Space rocket is made of two large halves of rigid insulating foam, with semi cylindrical lengths cut out of it for fit half of the rocket. The two halves are joined around the vehicle, and lightly bonded to each other to avoid falling off from wind or gravity. These halves are not only secured to the ground by better means, but also have a wide surface area to produce large amounts of drag. If the insulating foam should come loose from its ground harnesses, the drag force and inertia will push the sabot halves off and the rocket will take off the rest of the way uninsulated.

XVIII. Main Engine Start. Starting the Rocket Engine Fill/Drain valve locations for the propellant tanks are to be included with the plumbing as with most custom

rockets, but if the user may choose, these ports may be attached directly to the propellant tanks. Because the tank ends are not designed for maximum mass efficiency, the thickness of the aluminum and the flat faced external side makes it easy to attach ports on the bottom as well as the top. A concentric tank like in the Mark 2 has it’s Fill/Drain ports at an entirely different drainage point than for the main feed lines. This reduces the length and part count of the main feed lines, and lets the main valve be installed right up against the propellant tank.

The main valve assembly that feeds the propellants into the engine has to be actuated on and off on demand. This has been tested and proven with 3 pneumatic actuation methods. One method which can save mass and provide a substantial force for turning cold valves is an external piston actuator. The external main valve actuator is a powerful pneumatic piston mounted and aligned outside the rocket and attached to a bar that rotates the ball valve with lever action. Once the piston pushes the bar to turn the valve, the rocket can take off and detach from the external piston. Because the piston is left behind, it saves the weight of carrying the hardware with it. The nature of the detachable connection is a reliability risk since the connection can come loose at any time.

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To overcome the low reliability of the external main valve actuator, and internal main valve actuator can be used instead. This system has a pneumatic piston mounted on a bulkhead which pushes on the valve arm from the inside. This system requires gas to feed it from the vehicle or a detaching line from the outside. The downside of this system is the added weight of the piston and any valves used for the pneumatics. Since the valve does not detach, there is little risk of the connection coming loose, but a powerful piston becomes heavy.

A further simplified system is the rotary actuator, which can be a pneumatic or gear motor actuator. This system requires wide spacing between the valves, but is directly connected to valve stems instead of through an armature. The rotary actuator has a few drawbacks, such as the lack of a long lever advantage, but the system is small and simple. The proximity to the valves also make these actuators vulnerable to cold from the valves.

An internal system is necessary for a second stage vehicle, so a second stage Mark 2 can use an electric rotary actuator to turn the valve causing its engine to start and the vehicle can take off from a special barstool stage adapter.

The ability to launch as a multistage vehicle is important for the MLVA. Most western rockets have multiple steps involved with vehicle staging. With the Saturn V, the lower stage cuts off, and then the interstage, which joins both the upper and lower stage, detaches from both ends, separating the two vehicles. As the upper stage is separate but in freefall, a series of ullage motors fire to bring the stored propellant back to the bottom by accelerating the upper stage forward. Now that the propellants are at the bottom again, the second stage engine can fire, while the lower stage is far enough away that it won't reflect the thrust back at the upper stage. This kind of system is fairly complex, which goes against the design philosophy of the MLVA. Russian rockets like the Soyuz does staging much differently, with a “barstool” or open air cage as an interstage. The staging sequence is also much simplified, with the second stage firing while still attached to the first. The thrust is then deflected by a cone that diverts the combustion gasses outside through the open cage. If the second stage is fired while the first is still burning, it eliminates the need for an ullage motor, and the entire staging sequence can be reduced to a matter of firing up the second stage motor and releasing the second stage. An MLVA rocket is well suited for a soyuz style staging mechanism, and an MLVA rocket can be fitted with a conical adapter module and a body tube can be modified to make the cage, while a piston adapter at the bottom of the sustainer can mount the two pieces together with shear pins. The result is an exceedingly simple staging mechanism that makes multistage vehicles much more achievable. The holes created by the barstool can also be covered lightly such that they blow away at engine start.

Figure 18.1. Barstool Stage adapter for multistage rockets. The second engine is hot fired while mounted in a piston fit with the stage adapter and held in place with additional nylon shear pins. The rocket thrust pushes the second stage away from the first while still joined, and the combustion gases escape out the side.

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XIX. Payloads and Modules. The Purpose of the Rocket The same radial bolt pattern used in the rest of the rocket can be used to mount other miscellaneous

modules instead of dedicating a single payload bay at the nosecone/fairing. Just like how the tank ends and bulkheads are mounted, additional modules may be bolted within a body tube or even join two sections together. An example of when a module would want to join two sections is for observation or cold gas thruster arrays, where access to the outside of the rocket is necessary, but where the user does not want to drill out holes in a body tube for it, or if it needs to be flush with the outside or even protruding.

For generic payload mounting, a payload may be mounted in from the side of a rocket at any point using a module not unlike the cubesat adapter shown in Figure 19.1. For double length cubesats and similarly sized nanosats, the 8” diameter architecture is a perfectly adequate housing, and includes extra space for insulation or for lines and wiring in the vehicle to be diverted around the cubesat. In terms of ease of manufacture, the payload bay may be made from readily available ¼” thick 8” tubing cut to the required length. The two ends may be turned and radial holes drilled to create the standard piston bolt pattern. The payload housing can then be welded or otherwise bonded to the inside wall and the appropriate external access milled into the side. The radial pattern also lets this payload bay be oriented in any direction.

The payload mounting can apply to other experimental modules, such as a cold gas Reaction Control module in Figure 19.2. An example of one such module is shown below. This module takes advantage of a commonly available 8” tube with 1⁄2” walls. The piston ends and radial bolt pattern are turned and milled into place just like with the payload module. With the metal already in place, the cold gas thruster nozzle may be milled into shape using a tapered reamer at four equidistant locations to create the locations of enough thrusters for pitch and yaw authority. Different patterns and designs allow the module to achieve different thrust and authorities. With the modular rocket architecture, this RCS module may be installed or hot swapped into an available position on the rocket, which also means it can be removed or updated at will without having to disassemble any more of the vehicle.

Figure 19.1. Cubesat Payload bay with dummy payload. Figure 19.2. Basic Cold Gas Thruster Array in MLVA module.

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XX. Internal Access and Integration Vehicle integration and maintenance is easy to neglect when designing a vehicle, but it is the moment

where everything comes together, literally and figuratively. Before being assembled, a rocket will be in several pieces, and each of these pieces need to be aligned and slid into place. The MLVA is horizontally integrated, with roller trunnions used to rotate and slide parts of the vehicle around. These trunnions are standardized blocks with positioned upwards facing wheels and can support multiple sizes of rockets as shown in figure 20.1.These trunnions have two points of contact, but their unmarking polyurethane wheels do not scratch or damage the surface. Because wheels are used, the rocket is always supported by radial loads and the rocket body itself can be rotated for easier access and manipulation. Each rocket section needs two of these trunnions to support it at each end. When the rocket sections need to be integrated, they are slid together, and the trunnions keep both at the same level.

Figure 20.1. Roller trunnion. Support and contact points for 6”OD and 12”OD rockets. The wheels are polyurethane caster wheels, and the pattern can be mounted on independent blocks or in a longer dedicated table.

Internal access of the vehicle can weaken the strength of the body tube if access hatches are cut out, so the

best way to obtain internal access is to remove the tube whole. A slide-on body tube works by sliding over the internals and attaching at the tube ends. Because of the modular architecture of each section and subsystem, removing and integrating these parts as well as the body tubes themselves is simplified.

XXI. Fin Construction and Attachment A rocket of the smaller scales intended for the current MLVA has several options for fin stabilization to suit

the circumstances. The aluminum body tube is where the fins ultimately attach, and is capable of being milled to create fin slots if necessary. Aluminum fins are fairly heavy for fins, but have a significant advantage in durability at high speed flight over composite fins. If the user is willing to support more mass, an aluminum fin will keep it’s edge well in handling and can withstand high temperatures. A specific feature of aluminum fins is that they may be welded or epoxy cold-welded directly to the aluminum body. With fin slots, the aluminum fins can be welded on both the inner and outer sides, providing a stronger and a longer bond than a butt weld. Additional reinforcement can be added to improve the cantilevered connection. If the rocket is designed with a fairly simple lower section, the fin can is easy to swap or replace. The aluminum is not without sacrifice. Not only will aluminum fins be heavier, but the weight will be at the very bottom of the rocket, which shifts the CG backwards and the required fin area up.

Besides metal, fins may be made of composites such as G10/G11 or CF plates. The difficulty with composites is that they are not a uniform isotropic material, and the integrity of the plate is entirely dependant on the integrity of the binder, which is usually epoxy. At high speed flight, the friction of leading edge contact will heat up the fins considerably. A heated composite introduces several weaknesses to the material. At a certain temperature,

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cured resin will approach or surpass a glass transition temperature Tg where the solid cured resin will become softer and more pliable. A fair way to understand this transition is with rubber, like those on cars or o-rings. Though tires will continue to get softer as they heat up and harder when they cool, at very cold temperatures, rubbers will become very hard and brittle, and at a much faster rate than before. This characteristic makes elastomers a poor choice for cryogenic conditions, whereas the consistently rigid ptfe will perform as expected. Epoxy can have a fairly low glass transition temperature, with a Tg of 139 °F for ordinary WEST SYSTEM 105 to about 450°F for a high performance EPON resins. Regardless of the resin, temperature will make the material less stiff and more susceptible to flutter. At even higher temperatures, the epoxy will decompose or burn before the fibers do. The weakening of the resin has a direct influence on the strength of the composite since the resin is the binder that stiffens the fibers that act as the stiffening element for the otherwise loose and flexible fibers. Because of this, composite fins cannot be expected to perform as well at high temperatures and high velocity than it would at low velocity.

Composite materials have several other failure modes to contend with, but delamination is one that becomes problematic at the leading edge where the plys may be exposed and peeled off or away. Should a layer delaminate, the increased flutter and surface area drag can destabilize the vehicle and create a chain reaction in the vehicle. This risk as well as those mentioned before should be considered when choosing what fins to use, and the improved mass fraction may not always be preferred over the increased risk of total loss of vehicle, particularly for prototype vehicles.

Attaching composite fins is performed using a different procedure than metal fins. Though with fins slots different material fins can be bonded, the strength of the bond may not be sufficient for some applications. Epoxy cannot bond to composites as well as it would have before both were cured, so the same bond strength as welds cannot be expected. Bonds to fiber layers also only occur with the outermost layer, so a layer may just shear off a layer. Instead of a slotted tube, fins may also be integrated as a fin can. Composite fin cans can be slid on or radially attached after being segregated. The strongest bond of a fin to a composite fin can is during the cure cycle, so a skin made using the Newman method can be left to gel, and the fins attached before finishing it’s cure in the oven. Any fillets added at this point will be well bonded to the fin can even if not as much to the fin itself.

Acknowledgements The authors would like to thank the larger amateur rocketry community for the relatively free sharing of

knowledge and experience which was instrumental in the assembly of the architecture outlined here. Watching other rockets succeed or fail is an essential process in furthering our understanding of rocket dynamics at a small scale, and it is upon the sacrifices made and the tears shed that we may be able to understand what went wrong and call upon these experiences in preventing them from happening again.

Special acknowledgements to the Friends of Amatuer Rocketry (FAR) and it’s members including but not limited to Mark Holthaus, John Newman, and Rick Maschek. Most of the manufacturing methods and concepts were developed or conceived in cooperation with Mark Holthaus, to whom the rockets of the MLVA system are named after. The composites manufacturing methods and engine manufacturing methods were made possible by the efforts and experiences of John Newman.

Derek Honkawa would like to acknowledge the contributions of the faculty and staff of California State University Long Beach who made the development and funding of the rockets possible, in particular the Dean Forouzan Golshani, Co-author Dr. Mahdi Yoozbashizadeh, Dr. Praveen Shankar, Dr. Eric Besnard, Filipe Coehlo, Mike Fritz, and Joe Wardell.

The efforts of different student rocket groups is a valuable resource, and the student rocket teams of CSULB, SDSU, and UCLA have made great progress with limited resources. The contributions of the National College Resource Foundation and the Air Force make these efforts possible.

The construction of the Mark Space Rocket was made possible by the generous donations and contributions of Foster Stanback.

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