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IAC 18,c4,5,10,x45214
1.
69th International Astronautical Congress
(IAC), Bremen, Germany, 1-5 October 2018. Copyright ©2018 by the International Astronautical Federation (IAF). All rights reserved. IAC-18-F1.2.3 Page 1 of 17 IAC-18,C4,5,10,x45214 Liquid Rocket Engine Design for Additive Manufacturing Jan Fessla *, Hanmo Shenb , Nihar Patelb , Tai Wei Chena , Suyash Ghirnikarb , Martin Van Den Berghec a Department of Astronautical Engineering, University of Southern California, 854 Downey Way, Los Angeles, California, 90089, fessl@usc.edu b Department of Aerospace and Mechanical Engineering, University of Southern California, 854 Downey Way, Los Angeles, California, 90089, c Department of Earth Sciences, University of Southern California, 3651 Trousdale pkwy, Los Angeles, California, 90089. * Corresponding Author Abstract This paper discusses the processes involved in the additive manufacturing of a regenerative and film-cooled liquid rocket engine with a thrust of 10 kN using Inconel 718, while detailing validation techniques. A description of the objectives and design constraints provide the context and motivations. Computational Fluid Dynamics (CFD) models were developed and provided the expected pressure and thermal regimes under regenerative and film cooling. Additionally, Finite Element (FE) models were used to predict the capabilities of the engine structure. A description of 3D printing methods highlights the benefits and limitations of the technology, specifically the influence the design of liquid rocket engines. A pintle injector is used, printed as a separate, easily removable and replaceable component. Issues related to overhangs, surface roughness, and shrinkage; aspects related to post-print processing and the need to minimize machining are discussed. Results from the CT scans of the engine and its components are presented. The paper also outlines the series of tests that will be performed on this engine to verify its performance and provide design basis for future works This engine will be used to power the reusable flight vehicle that is under development at the Kyushu Institute of Technology in Japan. The student-led Liquid Propulsion Laboratory at the University of Southern California is responsible for the work detailed below. Keywords: Liquid Propulsion, Additive Manufacturing, Cooling Channels, Film Cooling, Pintle Nomenclature h = Heat Convection Coefficient Cp = Specific Heat μ = Dynamic Viscosity ρ = Density c* = Characteristic Velocity 𝑚̇ = Mass Flow Rate λ = Vaporization Heat Pr = Prandtl Number Re = Reynolds Number Nu = Nusselt Number T = Temperature P = Pressure R = Chamber Radius D = Chamber Diameter A = Chamber Cross-section Area p = Chamber Circumference S = Area Covered Gaseous Cooling Film x = Distance Downstream the Coolant Injection E = Young's Modulus α = Thermal Expansion Coefficient ν = Poisson's Ratio σ = Stress Subscripts aw = Adiabatic Wall Condition hw = Hot-side Wall Condition cw = Cold-side Wall Condition co = Coolant Property c = Property of Combustion Gas 0 = Stagnation Condition l = Property of Liquid Coolant g = Property of Gaseous Coolant s = Saturation Condition I = Initial Condition t = Throat Condition i = Inner Side Chamber Geometry o = Outer Side Chamber Geometry eff = Effective Parameter Θ = Tangential Component r = Radial Component x = Axial Component Acronyms/Abbreviations AM Additive Manufacturing CA Center Annulus CAD Computer Aided Drawings CFD Computational Fluid Dynamics LOX Liquid Oxygen
2.
69th International Astronautical Congress
(IAC), Bremen, Germany, 1-5 October 2018. Copyright ©2018 by the International Astronautical Federation (IAF). All rights reserved. IAC-18-F1.2.3 Page 2 of 17 REPDA Rocket Preliminary Design and Analysis TMR Total Momentum Ratio WFTS Water Flow Test Stand WIRES Winged Reusable Sounding Rocket Vehicle 1. Motivation This rocket engine, codenamed Balerion (Fig. 1), was designed by the Liquid Propulsion Laboratory (LPL) at the University of Southern California to meet the requirements of the reusable flight vehicle called the Winged Reusable Sounding Rocket #13 (WIRES #13) (Fig.2), currently under development at the Kyushu Institute of Technology by the Kyutech Space Club. To meet the requirements, two engines with a total thrust capacity of 20 kN and a burn time of 25 seconds have been proposed. LPL is responsible for the development of the propulsion systems which includes the engines and propellant feed lines. This paper will focus on the analysis, design, manufacturing, and testing of the engine. Attention and deliberation were focused on the use of additive manufacturing to minimize the number of components and gain a deeper understanding of the technology that is sweeping the industry. Table 1. Engine parameters Propellants LOX/Kerosene Chamber Pressure 375 psi O/F ratio 1.5 Chamber Temperature 2464 K Thrust 10 kN Specific Impulse 241 sec Total Mass Flow Rate 4.25 kg/s Fig. 1. CAD model of Balerion engine 2. Theoretical Background In the preliminary design and 1-Dimensional (1D) analysis phase of Balerion, an in-house Rocket Preliminary Design and Analysis (REPDA) MATLAB code was used to determine all aspects of the thrust chamber and nozzle design. Fig. 2 CAD model of WIRES #13 2.1 Propellant and Material Selection During the conceptual design phase, a propellant trade study was conducted on two fuel types – Unleaded kerosene (Jet-A) and isopropyl alcohol (IPA). Ultimately, Jet-A was selected mainly because it has more chemical and fluid properties readily available. Due to the engine utilizing additive manufacturing material selection options were very limited. Materials that are available and applicable for a rocket engine include high-strength nickel-base superalloy Inconel 718 and aluminium alloys, in particular AlSi10Mg. The material properties of interest were thermal conductivity for cooling performance and yield stress for structural integrity. The comparison is shown in Table 2. Table 2. Comparison of AlSi10Mg and Inconel 718 Property AlSi10Mg Inconel 718 Thermal Conductivity 173 W/(mK) 11 W/(mK) Yield Stress 228 MPa 1241 MPa The deciding factors between the two were the low yield stress of AlSi10Mg and the higher availability of Inconel 718 for printing. It was determined that AlSi10Mg was not a viable option when it came to withstanding the operational pressures during testing and lacked detailed data sheets for its temperature capabilities. However, Inconel 718 has more readily available information regarding its properties. With slight material modification (heat treatment) can handle the required operational conditions of the engine. Table 3 shows the properties of Inconel 718 after undergoing AMS 5662 heat treatment. Table 3. AMS 5662 properties of Inconel 718. Physical properties are at 992 K Melting Point Heat-treated 1571 K 1650 K Thermal Conductivity 22 W/m/K Yield Stress 1010±50 MPa Young’s Modulus 2.25×105 MPa Thermal Expansion Coefficient 1.51×10-5 K-1 Poisson Ratio 0.283 2.2 Engine Sizing using REPDA Software For propellant analysis, the relationship between O/F ratios at different chamber pressures and chamber
3.
69th International Astronautical Congress
(IAC), Bremen, Germany, 1-5 October 2018. Copyright ©2018 by the International Astronautical Federation (IAF). All rights reserved. IAC-18-F1.2.3 Page 3 of 17 temperature or specific impulse was given by REPDA. REPDA was specifically developed to size a rocket engine chamber and nozzle based on its thermal performance requirements. It builds upon the freely available NASA Chemical Equilibrium with Applications (CEA) software, and the NIST Reference Fluid Thermodynamic and Transport Properties (REFPROP) database. Examples of the REPDA outputs include chamber pressure, propellant mixing ratio, thrust chamber geometry, cooling channel geometry, along with many others. 2.3 Thermal Analysis To determine the optimal thermal properties of the engine, various iterations of differing chamber pressures, O/F ratios, and propellant mixtures were used. REPDA made these iterations quick and easy due to its user-friendly GUI system as seen in Figure 3. Fig. 3. Showing Propellants Input to REPDA 2.3.1 Regenerative Cooling Using REPDA, inputting the desired rib thickness, hot-wall thickness, minimum channel width, and channel depth provides the required number of channels needed to keep the temperature under the materials melting point and thermal stress below the materials yielding point. Additionally, printing surface roughness can be inputted to determine effects on cooling performance. This can be seen in Figure 4. The governing equation used in this 1D analysis is the Bartz equation (1). ℎℎ𝑤 = 0.026 𝐶𝑝 𝑐 𝑃𝑟𝑐 0.6 ( 𝜇 𝑐 𝐷𝑡 ) 0.2 ( 𝑃𝑐,0 𝑐∗ ) 0.8 ( 𝐴 𝑡 𝐴 ) 0.9 𝑆 (1) Fig. 4. Input in REPDA for regenerative cooling calculation While seeking the optimal design point of the regenerative cooling system, the convection coefficient correlation specifically dedicated to RP-1 at supercritical pressure [1] was embedded into the REPDA subroutines and is referenced as Eq. (2). In terms of the thermophysical properties RP-1, a believable data source from NIST [2] was also merged into a spreadsheet accessible through the code. Nu 𝑐𝑜 = 0.012 𝑅𝑒 𝑐𝑜 0.879 𝑃𝑟𝑐𝑜 0.4 (1+ 2Dℎ 𝐿 ) (2) The temperature profiles calculated and shown by REPDA inherently contain ±20% of uncertainty introduced by the Nu number correlation which suggested another cooling method be used to assist regenerative cooling and help promote full engine reusability. This introduced the need for film-cooling to the engine discussed in Section 2.3.2 2.3.2 Film Cooling To model the liquid film-cooling and start the parametric analysis, a film-cooling model developed by Stechman [3] was implemented. This model was determined to be applicable solely for small rocket engine designs. Both liquid and gaseous phases of the film-cooling were taken into consideration. All the equations were based on the following assumptions. First, the gaseous coolant film exhibits no mixing or chemical reaction with the main combustion fluids. Second, the coolant film temperature profile does not change rapidly as it moves downstream. Third, the gradients across the coolant film are small. Finally, heat transfer to the wall is negligible compared to the heat transfer from the main combustion fluids. The length of the liquid-cooled region was calculated by Eq. (3), and the heat convection coefficient for the liquid phase was calculated by Eq. (4). 𝐿 = 𝜂𝑚̇ 𝑙 𝐶𝑝 𝑙(𝑇𝑠−𝑇 𝐼) 𝑝ℎℎ𝑤(𝑇 𝑎𝑤−𝑇𝑠) + 𝜂𝑚̇ 𝑙 𝜆 𝑝ℎℎ𝑤(𝑇 𝑎𝑤−𝑇𝑠) (3) ℎ𝑙 = 0.0288 𝐶𝑝 𝑙 𝑃𝑟𝑙 0.667 𝜇 𝑙 0.2 𝑥0.2 ( 𝜂𝑚̇ 𝑙 𝜌 𝑙 𝑢 𝑐ℎℎ𝑤 𝑃𝑟𝑐 0.667 𝜋𝑅𝐶𝑝 𝑐 ) 0.4 (4) As for the gaseous film cooling, effective gas temperature Teff was introduced as an alternative to the adiabatic wall temperature. This temperature was calculated according to Eq. (5). The gaseous convection coefficient, as mentioned earlier, was concluded from Bartz equation, see Eq. (1). The film cooling inputs for REPDA are shown in Fig. 5. 𝑇 𝑎𝑤−𝑇 𝑒𝑓𝑓 𝑇 𝑎𝑤−𝑇𝑠 = 𝑒𝑥𝑝 ( −ℎℎ𝑤 𝑆 𝜙𝐶𝑝 𝑔 𝑚̇ 𝑔 ) (5)
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(IAC), Bremen, Germany, 1-5 October 2018. Copyright ©2018 by the International Astronautical Federation (IAF). All rights reserved. IAC-18-F1.2.3 Page 4 of 17 Fig. 5, REPDA inputs for film cooling 2.3.3 Thermal Stresses The last consideration of 1D analysis is running a thermal stress load study on the thrust chamber, which is caused by the temperature difference between the inner and outer surfaces of the wall. The thermal stress on the structure was modelled as a hollow cylinder being heated up from inside. There were three stress aspects considered: tangential, radial and axial. These formed the effective thermal stress model according to the Von-Mises theory. The calculation was executed using Eq. (6). 𝜎𝑣 = √𝜎 𝜃 2 + 𝜎𝑟 2 + 𝜎𝑥 2 − (𝜎 𝜃 𝜎𝑟 + 𝜎 𝜃 𝜎𝑥 + 𝜎𝑟 𝜎𝑥) (6) 2.4 Injector An injector is one of the key components in a liquid rocket engine as it determines how the propellants are delivered into the combustion chamber as well as their mixing and combustion efficiency. Among the various types of rocket engine injectors that exist, the Balerion engine uses the pintle as it has inherently strong combustion stability, reliability, design simplicity, and ease of manufacturing compared to the traditional showerhead and coaxial injectors. The pintle injector delivers one propellant radially through the pintle and the other propellant axially in the shape of an annular sheet through the injector body to impinge and form a uniform conical stream. Due to this spray cone formation, a recirculation zone is created right below the pintle tip, which acts as a mixer for unburnt propellant droplets, further improving the combustion efficiency and stability of the engine [4]. 2.4.1 Key Parameters The most important parameter for an injection system is the total momentum ratio (TMR), which is defined as the momentum ratio of the radial-to-axial propellant stream, seen in the following equation: 𝑇𝑀𝑅 = (𝑚̇ 𝑈) 𝑟 (𝑚̇ 𝑈) 𝑧 (7) It has been shown that pintle injectors with a TMR close to 1 provide the most optimal performance. Another key variable, known as blockage factor, is defined as the ratio of total hole diameters to the pintle circumference, with the equation as: 𝐵𝐹 = 𝑁𝑑0 𝜋𝑑 𝑝 (8) where N is the number of orifices at pintle tip, d0 is the orifice diameter, and dp is the pintle diameter. This value pairs with TMR, orifice size, and pintle geometry to affect the propellant mixing efficiency [5]. In addition, some recommended values of the pintle injector dimensions are given. For instance, the ratio of chamber to pintle diameter is recommended to be between 3 and 5, while the skip distance, defined as the axial distance between the annular ring exit and the radial hole, is suggested to be a 1 to 1 ratio with the pintle diameter [5]. 2.4.2 Pressure Drop Consideration Accurately predicting the pressure drop across the injector is crucial for the engine to perform nominally. It is recommended to have a pressure drop of at least 20% of the chamber pressure on both propellant sides in order to neutralize the possibility of chamber pressure oscillations. This would create combustion instability that can build up and severely damage the hardware if it were to backflow through the injector [6]. On the other hand, if the pressure drop is too high, it can place extra design limitations on the propellant tanks since the tanks need to be sized based on the internal pressure requirements. The higher the pressure requirement, the harder the tank will be to design and manufacture. Additionally, this would cause issues for the feed system that delivers the propellants because it is specifically sized to maximize performance while minimizing weight. 2.4.3 Injector Sizing Normally, when sizing a traditional injector, the discharge coefficient (Cd) value of the orifice type will be used along with the following equation to determine total orifice area at a given mass flow rate, propellant density and desired pressure drop: 𝐶 𝑑 = 𝑚̇ 𝜌𝑉̇ = 𝑚̇ 𝜌𝐴𝑢 = 𝑚̇ 𝜌𝐴√ 2∆𝑝 𝜌 = 𝑚̇ 𝜌𝐴√2𝜌∆𝑝 (9) However, since the typical Cd value of the pintle injector was not found in the initial sizing phase, a different model, mainly utilizing Bernoulli's principle, was developed to size the injector. For an injector to perform optimally, the TMR value needs to be as close to 1 as possible, which depends heavily on the propellant outlet velocities since the mass flow rates are fixed from initial engine sizing. Therefore, a MATLAB code separate from REPDA was developed to vary the number and size of the orifices at
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(IAC), Bremen, Germany, 1-5 October 2018. Copyright ©2018 by the International Astronautical Federation (IAF). All rights reserved. IAC-18-F1.2.3 Page 5 of 17 the pintle tip to alter the central propellant velocity flowing out from the orifices. Total orifice area at the pintle tip was taken into account by using Bernoulli’s equation (10) in an attempt to find the pressure drop across the orifices. Assuming the elevation is negligible, the pressure difference can be found from the change in area when flow exits from the orifice into the chamber. After that, the fuel outlet velocity can be found from the TMR equation (7) at a value close to 1, which determines the fuel side annular gap thickness as well as its pressure drop. 𝑣2 2 + 𝑔𝑧 + 𝑝 𝜌 = 𝑐𝑜𝑛𝑠𝑡𝑎𝑛𝑡 (10) 3. Design This section covers the basis upon which the engine was initially sized using results generated by REPDA and how these numbers were implemented into creating CAD model. 3.1.1 Printing and Manufacturing Constraints The major limitation to any design is its manufacturability. With additive manufacturing, the driving constraints were established by the volume of the printing instruments; the total working volume is 250×250×325 mm3 (9.85×9.85×12.8 in3 ). Other prominent design constraints would be material availability, minimal wall thicknesses (0.45 mm, 0.0177 in) and overhangs in the printing direction (≥ 45°) and critical considerations in design geometries. 3.2 Input Parameters The input for initial dimensions of the engine such as dimensions of regenerative cooling channels, outputs generated by REPDA were used. 3.2.1 Results using REPDA Software Fig. 6 indicates that the specific impulse of LOX/Jet-A propellant pair hits the theoretical maximal value at O/F ratio around 2.2, at which the chamber temperature also raises up to 3400 K. Given the potential cooling challenge on the thrust chamber, particularly with the low thermal conductivity of Inconel 718, an O/F ratio of 2.2 was not considered feasible, and eventually an O/F ratio of 1.5 was determined to be an optimal trade-off between specific impulse and chamber temperature. Chamber pressure selection was chosen according to the supply pressure of the propellant tanks inside the WIRES#013 vehicle, to be 375 psia. Fig. 6. REPDA results showing relation between O/F ratio and pressure to Specific impulse and combustion temperature. 3.2.2 Injector Sizing Considering cooling systems as well as overall integration purposes, it was decided early in the design stages to have Jet-A be the regenerative coolant entering the chamber down the annular ring, and LOX flow through the pintle, entering the chamber radially from the pintle tip. It was however discovered through CFD analyses during the iterative design process that the LOX pintle outlet geometry, while maintaining a TMR of 1, would not meet the required pressure drop of 20% greater than chamber pressure. As a result, an additional “throat” feature was added at the inlet of the pintle with the purpose of making the pressure regime in the LOX system meet requirements. This pintle pressure drop, as modelled in 3D simulations, was later proven to be inaccurate compared to empirical testing, underestimated the added effects of possible frictional loss and energy dissipation during fluid flow. While 3D modeling can take into account such parameters, quantifying and inputting them into such models is challenging as some of these effects result from the printing process, and are generally still poorly constrained. Nevertheless, CFD can still help size unknown orifice geometries in a design iteration process. 3.3 Balerion Development Engine Model Fig. 7 is the final model of the Balerion Development Engine designed using Siemens NX Unigraphics. It stands at 448.31 mm (17.65 in) tall with a max diameter of 218.44 mm (8.30 in) at the flange.
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(IAC), Bremen, Germany, 1-5 October 2018. Copyright ©2018 by the International Astronautical Federation (IAF). All rights reserved. IAC-18-F1.2.3 Page 6 of 17 Fig. 7. Balerion Development Engine Model In the initial concept design, the goal was to minimize the number of components to reduce complexity, thus improving reliability. This prompted us to have only three components in total, two for the injection system and one main engine body. However, early on it was discovered that most printers under service are restricted by printing volume, rendering the unibody engine architecture unattainable. This led to the decision to split the main body into two components and join them with flange. One component became the upper chamber body, while the second component became the nozzle with the convergent and divergent sections. The two remaining components form the pintle injection system. The pintle was designed to be interchangeable so that a number of orifice configurations, skip distances and pressure drops can be tested during cold flows to select the most optimal empirical based design model for engine performance. The center annulus serves as a connecting adaptor for the pintle and the chamber. Once inserted into the chamber body, it forms an outer annular ring for the fuel to be discharged into the chamber to form a propellant mixture after it travels through the regenerative channels. The pintle works on the concept of radial injection while the outer annulus works on an axial injection basis. The interaction of an axial fuel stream combined with a radial outlet creates a conical spray pattern. More design changes were made based on analysis, manufacturability considerations, legacy engine data, and available resources. For more information on the specific design features of the engine, refer to Section 3.4. 3.4 Design Features As previously mentioned, many different considerations were taken before reaching the model shown in Fig. 7. The most critical considerations are expounded upon in the following subsections. 3.4.1 Regenerative Cooling Channels For the regenerative channels, the driving decision for the geometry was based on the limitations of the printer (For more information on the printer, refer to Section 5.1.1.). The fluid acting as the coolant for Balerion is Jet-A. The channels are rectangular slots (shown in Fig. 8) that run from the divergent section of the nozzle up into the injector. The preliminary wall thickness of the chamber separating the combustion gases and the flowing coolant was 1 mm but was altered due to data gathered on printing capabilities and referencing existing printed components. The minimum wall thickness is extremely case sensitive, and for Balerion, the smallest possibility was around 0.3-0.4 mm. However, to ensure that the print does not fail, 0.6 mm was the decided minimum wall thickness. REPDA was used to iterate the channel ribs and the thickness ended up decreasing from 1 mm to 0.6 mm. This automatically increased the number of channels based on the engine cross-section which allows for better cooling of the engine body. The number of channels increased from 124 to 150. With a constant rib thickness, the channel size varies while traveling axially up the engine cross-section, thereby providing the highest cooling at the throat due to that region having the most restricted slot openings (0.65 mm). For a smaller channel cross-section, the fluid moves at a higher velocity which corresponds to having better cooling as more energy is dissipated away from the hot- wall into the coolant which is then injected back into the chamber. The only unchanged parameter was the channel depth, which remained constant at 1 mm. Refer to Table 4 for more information on channel sizes. Fig. 8. Regenerative Channels as shown in a chamber cross-section. Table 4. Final dimensions of regenerative cooling system Inner chamber wall thickness 0.6 mm Channels depth 1 mm Rib thickness 0.6 mm Channel maximum width 2.9 mm Channel minimum width 0.63 mm Number of channels 150 While maximizing the effectiveness of regenerative cooling through modelling and iterative design, it was still apparent that the thermal stress caused by the steep temperature gradients in the Inconel structure could
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(IAC), Bremen, Germany, 1-5 October 2018. Copyright ©2018 by the International Astronautical Federation (IAF). All rights reserved. IAC-18-F1.2.3 Page 7 of 17 potentially cause structural failures. REPDA analysis indicated that the peak stress occurs at the engine’s throat and is 130 ksi, which approaches Inconel’s yield stress of 146 ksi at 922 K. For this purpose, film cooling was added as a means to control the engine’s thermal and stress regime. 3.4.2 Film cooling The Film cooling orifices are located above the flange on the chamber side. The orifices serve as a thermal management system by injecting a pre- calculated amount of fuel into the chamber in a fuel rich setting. The orifices are 0.38 mm in diameter. The number of orifices is the same as the number of channels. 3.4.3 Sensors To fully characterize the empirical performance of the engine, sensors are placed in various places for measuring pressure and temperature (Table 5). Testing ranges from measuring the pressure drop on the LOX and regenerative cooling channel while visualizing the cone formation. LPL’s Water-Flow Test Stand (WFTS) was used for testing [7]. Fig. 9. Sensor placement in the injector & chamber The left cross-sectional view (Fig. 9) provides insight into how these sensors are placed to measure pressure and temperature in the most critical areas. On the top, three sensor ports are located in the center annulus. Two sensors measure pressure and the other measures temperature. The same architecture is used for the fuel side of the injector. The chamber has four pressure transducers which are located evenly around the circumference of the chamber and are used to measure the chamber pressure. The multiple sensors will help in determining pressure distribution in the chamber and help to detect combustion instabilities. The last group of sensors is located in the cavity at the end of the nozzle. Five sensors are used in total to fully characterize the flow properties in the nozzle cavity. Two pressure transducers are located symmetrically around the fuel inlet. The purpose of Table 5. Sensor configuration on development engine Position of Sensors Number of Sensors Purpose Pressure Transducers Nozzle exit 4 Coolant pressure Chamber 4 Chamber pressure Injector 3 Fuel & LOX pressure Thermocouples Nozzle cavity 2 Coolant temperature Injector 2 Fuel & LOX temperature Fig. 10. Sensor Placement in nozzle cavity these sensors is to measure flow distribution on both sides of the cavity to ensure they are even. The third pressure transducer is located on the side opposite to the fuel inlet. Fig.10 shows the variable cross-section manifold from the fuel inlet toward the other side of the nozzle. This is provided to keep the velocity of the fluid constant as it is distributed to channels. Two thermocouples are also located in the nozzle fuel manifold. The value from these sensors will be compared with the temperature readings downstream of the regenerative cooling channels to determine the cooling performance during the hot fire testing phase. 3.4.4 Constant Velocity Manifold The constant velocity manifold acts as a flow distributor to ensure that the fuel is traveling at the same rate axially through the regenerative channels of the engine. This is key for the ignition priming sequence because the fuel needs to be discharged proportionally as it will directly affect cone formation and thereby the engine performance. Reference Fig. 11 to see the cut away view of the constant velocity manifold.
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(IAC), Bremen, Germany, 1-5 October 2018. Copyright ©2018 by the International Astronautical Federation (IAF). All rights reserved. IAC-18-F1.2.3 Page 8 of 17 Fig. 11. Constant Velocity Manifold / Fuel Inlet 3.4.5 Fuel Inlet The fuel inlet (as seen in Fig. 11) is the latest major addition to the engine. It is also printed out of Inconel 718 and its purpose is to inject the fuel into the engine without having to drill directly into the nozzle. This component was developed and integrated to address the concern of clogging the channels from metal dust and shavings during the subtractive machining phase. In the preliminary design, there was a straight thread feature that would connect directly with a flex hose which would increase possibility of clogging the channels. With the addition of the fuel inlet there is lower chance of machining failure. 3.4.6 Flange There were five major aspect to consider during the design of the engine flange. The number of bolts, location, seal, coolant transfer cavity, and alignment (Fig. 12). There are a total of 18 bolts on the flange that allow to reduce the preload per bolt and to gain a higher structural factor of safety. The preload per bolt required to ensure o-ring compression is 6357 N, providing a factor of safety of 4.2. The thermal and stress load on the flange was a primary concern during design. Therefore, the flange was placed right below the film cooling orifices in order to maximize cooling effects on both the flange and throat. For more information about the film cooling holes, refer to Section 3.4.2. After the location had been decided, sealing the fluid in the regenerative channel while keeping the hot gases inside the chamber was the next challenge. To solve this, two metal o-rings were integrated on each side of the coolant transfer cavity to seal off the various fluid compartments (combustion and coolant). O-rings made of Inconel 718 with a silver coating were chosen due to their thermal capabilities. The fluid transfer cavity through the flange was used as a way to connect the 150 regenerative cooling channels between the chamber and nozzle while avoiding alignment issues. The fluid transfer cavity serves as a ribless zone that pools the fluid from one component to the other without needing a specific orientation. Two chamfers were added onto the sides of the cavity as an extra safety feature to ensure that after torquing down all the bolts, the channels will not have any flat overhangs that would cause a heavy pressure drop during operation. The need for an alignment feature was driven by the 1 mm opening of the fluid transfer cavity. With such a small opening, it is important to ensure that they align almost exacty to not affect engine performance. To achieve this, precision dowel pins are used on two of the eighteen bolt holes. The two bolts are oversized to 10 mm to ensure that they cannot be placed into any other bolt hole. Fig. 12 Engine Flange cross section 3.4.7 Pintle and Center Annulus The Pintle and Center Annulus (CA) as seen in Fig. 13 are the main components of the injector. The key design considerations were to make the pintle modular to test various development models. Initial design calculations were made to size the baseline model. For more details on sizing, refer to Section 2.4. Through the use of the WFTS, the pintle will undergo various tests to determine the percent error of the design values, and then iterated to achieve an optimal design. Details of the tests can be found in Section 6.1.1. The modularity of the pintle comes from its interchangeability with the CA. As long as the threaded section remains constant, various pintles with modified skip distances, orifice diameters, or flow diverters can be tested to empirically characterize the design. There is a Teflon o-ring seal in between the pintle and CA and also between the CA and chamber to prevent any unwanted leakage of combustion fluids. In order to switch out pintles with the CA, tooling features were printed into both components. For the CA, a hex feature was sized to ensure it can be placed within a vice grip for fastening and worked on with a standard wrench. The pintle bottom hex socket was designed to be used with a standard torque wrench to achieve the required preload which would be determined via testing. 3.4.8 Fitting Selection All fittings consist of an MS Boss/SAE configuration on the side that connects into the engine. The pressure transducer and temperature sensor ports convert to Swagelok™ while the
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(IAC), Bremen, Germany, 1-5 October 2018. Copyright ©2018 by the International Astronautical Federation (IAF). All rights reserved. IAC-18-F1.2.3 Page 9 of 17 Fig. 13 Pintle/Center Annulus fuel and LOX inlet ports to AN fitting, to comply with the requirements of the test setup. The reason for using MS Boss fittings is to have a pre-integrated sealing device on the fittings. This will prevent the hassle of performing sealing studies and procuring special sealing devices for the fittings. MS Boss fittings are highly adaptable to different kinds of fittings and come in a wide range of sizes. 3.4.9 Integration For empirical testing of the engine at its design hot fire conditions, an integration tool was required. The tool would need to align and connect the engine to the test setup and act as a thrust transfer mechanism. The main failure mechanism for the part would be buckling. The tool would also need to have enough clearance to pass flex hoses and other plumbing connections while providing easy access for assembly and disassembly. The tool would have a bolt pattern that would mate with adequate accuracy on to the engine. The bolt pattern on the engine is shown in Fig. 14. There are 9 bolts in total which are designed to handle the thrust and the moments from the engine. Fig. 14. Integration Bolt Pattern 4. Numerical Simulation To confirm results obtained by REPDA, 3D CFD analysis were performed on all engine components using Ansys Fluent software and supercomputer services provided graciously by Nimbix. These 3D CFD analyses were used to understand the thermal regime of Fig. 15. CFD temperature profiles of the regenerative system (without film cooling). The curved plane shows heat flux along the throat wall, the three straight planes show cross-sections of temperature across the combustion fluids, the Inconel and the coolant. the solid engine structures as well as coolant and combustion fluids. However, film cooling was not included in thermal simulations du to the complexity and computing requirements associated with such complex 3D analyses. 4.1 Thermal Results The thermal results without considering film cooling described a very steep temperature gradient through the throat wall at about 100 K per 0.1 mm, which was also predicted by REPDA. Furthermore, analysis showed that a thin veneer of the coolant was in contact with the channel. The channel temperature was predicted to be 700 K, but the bulk coolant temperature never exceeds 460 K (Fig. 15 and 16). This predicts that at ~500 psia the coolant will not boil. CFD analysis could help predict the effectiveness of regenerative cooling at the flange, an otherwise complex structure which 1D analysis does not account for. Despite the coolant being diverted away from the hot wall through a flange manifold, the highest temperature at the flange (without film cooling) is around 1150 K. Fig. 16. Temperature profile of the engine hot wall along the throat and flange, without film cooling.
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(IAC), Bremen, Germany, 1-5 October 2018. Copyright ©2018 by the International Astronautical Federation (IAF). All rights reserved. IAC-18-F1.2.3 Page 10 of 17 Fig. 17. CFD model of LOX pressure drop through the pintle injector. 4.2 Fluid Results To ensure that combustion products do not backflow into the feed system, CFD helped define an improved pintle geometry that would provide a 19% pressure drop (Fig. 17) for the LOX flow. This was achieved by including a throat feature upstream to achieve an effective LOX pressure regime while maintaining the desired TMR. 3D CFD analyses has helped in the design of an appropriate injector capable of providing a 22% pressure drop for the kerosene entering the chamber, while similarly providing appropriate relative mass flow rates between the outer annulus and film cooling (86% and 14% of total kerosene mass flow rate respectively) (Fig. 18). Along with helping narrow down design geometries, CFD proved effective at providing test system requirements to meet propellant flow criteria. The pressure drop in a full-length cooling channel under nominal hot fire conditions was quantified at ~100 psia (Fig. 19). These results illuminated benchmark requirements for feed system to maintain the desired engine performance under hot fire conditions. Fig. 18 Pressure drop through the outer annulus for fuel side of injector Fig. 19. Nominal pressure drop along the entire length of the regenerative cooling channel. 5. Manufacturing This section describes the basic features of the additive manufacturing (AM) post-printing process and inspection of printed components 5.1 Additive Manufacturing Selective Laser Sintering (SLS) technology printers have reached a stage where they are capable of printing complex structures using a wide variety of materials. This manufacturing process may not be cost effective but saves a large amount of time, effort and material. The overall benefits not only include resource savings but also allows the organization to research and validate additively manufactured structures for service. AM does come with inherent disadvantages, which lead to design complications. These problems were interpreted during the preliminary design and helped save excessive changes to CAD during the final stages. 5.1.1 Printer Specification The engine was printed on an EOS M290 machine. Industrial 3D printers are limited on build volume and accuracy so as to maintain a large amount of reliability and precision. Table 6. Printer Specification [13] Parameter Value Units Build Plate Size 250×250×325 mm Resolution ±40 μm Roughness ±15 μm Min. Wall Thickness 0.3 ~ 0.5 mm Laser Speed 7 m/s Laser Power 400 W Focus Diameter 100 μm 5.1.2 Print Inputs The most common AM constraint is the angle at which an object can lie when unsupported from beneath, an issue of overhang. All the surfaces that do not have supports underneath them would need be placed at an
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(IAC), Bremen, Germany, 1-5 October 2018. Copyright ©2018 by the International Astronautical Federation (IAF). All rights reserved. IAC-18-F1.2.3 Page 11 of 17 angle greater than 45º to the horizontal. With any angle less than 45º to the horizontal, the material would need to be supported by a mesh or solid structure. This severely restricts the geometry that you can have given that surfaces need to have a good surface finish or are thin wall sections that cannot be machined to remove supports. The flange section of the chamber was provided with a 45º solid support section because the chamber was printed flange up. The reasons for printing the chamber flange up are mentioned in the upcoming sections. The nozzle, however, was printed flange down and by doing this no sections were at an angle less than the recommended limit. The center annulus was also supported via a 45º ribbed support. Another consideration for manufacturing is the print direction. This refers to the orientation of the parts and how they will print as the deposition goes on. The parts can be printed at various angles and do not need to have a flat plane on the build plate. This is particularly helpful as the print direction dictates the strength of the parts. In the case of Balerion, the chamber was printed flange down to prevent the need of a support structure on the interior of the chamber because the thin walls of the section would not have been able to withstand the stress. This meant that a solid support was needed to provide stability for the flange. Fig. 20 Printed on parts on build platforms Studies have shown the a relationship between print orientation and mechanical properties [14]. The placement of parts on the build plate largely affects the success rate of the print. This means that parts must be properly spaced to avoid a print failure which has a very high monetary impact. The pintle was printed in its functional orientation with the injection holes closer to the build plate. The same applies to the center annulus, which was printed with the pintle mating face on the build plate. These directions were all in the +90º or -90º directions which ensured consistent mechanical properties. The print success rate is also highly affected by the amount and direction of heat transfer that can occur on the build plate. This is the reason why no two parts were placed on the same build plate. The pintle and the nozzle were printed on the same plate initially and it was noticed that the surface quality of the pintle was poorer than that of the nozzle. A new plate was configured to accommodate the small parts such as the center annulus, the pintle and fuel inlet on the condition that they could be spaced at a sufficiently large spacing. The files were provided to the print company in the .STL file format. The parts were modeled in NX 11, which allows users to set the resolution on the generated file. To ensure maximum resolution and accuracy from the printer the file resolution was set to 0.0001. This file format is capable of converting a solid body file into a mesh-based file for the printer to interpret. 5.2 Post–Printing Inspection In addition to AM, several post-processing technologies, including surface and heat treatment are required for the final product to accomplish the required thermal, mechanical and surface properties. Inconel-718 when heat treated as per AMS 5664 develops a hardness close to 47 HRC. For the main design considerations, some of the thermal and mechanical properties of AMS 5664 are listed in Table 3. To verify that all critical features were printed as intended various techniques were used. They are enumerated in the following sections. 5.2.1 Powder Removal The AM process leaves a large amount of unbound powder in the channels and closed sections which needs to be removed. There are multiple methods which are recommended for powder removal. For this specific application, the parts were first placed on a shaker table and close to 80 % of the loose powder was removed by vibrational means. The shaker table is initially set to a low-frequency setting and then the frequency increases to remove more powder. After this, the parts underwent a compressed air blowdown to remove any stray powder stuck in any low points of the system. The parts were then ultrasonically cleaned in an industrial grade ultrasonic cleaner for 1 hour each. The parts were CT scanned before and after the ultrasonic cleaning to determine the effectiveness of the procedure. 5.2.2 Heat Treatment The printed parts do not have the surface properties that are recommended for liquid rocket engines. The heat treatment process was composed of solution treatment and annealing at the recommended temperatures and times. Before heat treatment, the product was closer to a light gray color; after heat treatment, a discoloration of the parts to a blue-gray tinge was noted. The heat treatment process is done before part removal to prevent excess distortion and warpage of parts during the heat treatment process. This
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(IAC), Bremen, Germany, 1-5 October 2018. Copyright ©2018 by the International Astronautical Federation (IAF). All rights reserved. IAC-18-F1.2.3 Page 12 of 17 Fig. 21. Regen Channels at the throat also ensures that any small particles of powder that are still present in the channel do not affect the Wire EDM process. 5.2.3 Part Removal The smaller parts were separated from the build plate using a band saw. The more critical and bigger parts such as the chamber and the nozzle were separated from the build plates using Wire Electrical Discharge Machining (EDM) with 0.1 mm tolerance. EDM was used due to the large surface area involved along with the high risk of blade breakage if a band saw was used. 5.2.4 CT scan The internal structure was inspected by using Computed Tomography (CT) technology. The primary dimensions for inspection were that of the regeneratively cooled channels. Multiple planes were created at inspection locations and measurements were taken at those locations. The width of the channels varies from the end of the nozzle to the top of the chamber. The width of the ribs, however, is constant. On Fig.22 the dimensions of one of the regenerative Fig.22. Regen Channels Inlet Fig. 23. Planes for channel inspection on the chamber cooling channels inlet is shown. The average width was determined to be 1.87 mm. The design value was 1.8 mm. The difference in dimension is attributed to shrinkage. The channels were then inspected at the throat. The dimension of the channels is the smallest at the throat due to its relatively small diameter (Fig. 21). Fig. 24 Detail view on side of the chamber channels Table 7. CAD/CT dimensions comparison (mm) Location Parameter CAD CT Nozzle base Width 1.8 1.87 Throat Width 0.67 0.63 Chamber top view Width 2.62 2.34 Chamber top view Depth 1 1.01 Chamber side view Width 2.62 2.35 Chamber side view Thickness 0.6 0.82 Pintle top view Diameter 1.8 1.73 Pintle side view Diameter 1.8 1.59
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(IAC), Bremen, Germany, 1-5 October 2018. Copyright ©2018 by the International Astronautical Federation (IAF). All rights reserved. IAC-18-F1.2.3 Page 13 of 17 Fig. 25. Top view of regen channels in the chamber The dimensions of the regenerative channels on the straight part of the chamber were also measured. Two planes one on the top view and the other on the side view (Fig. 23 Blue & Green plane) were used to measure the width of the channels. The details views are shown in Fig. 24 and 25. The last part that was inspected was the pintle injector. The most critical dimension being the orifice diameters around the pintle tip. They were inspected from (Fig. 26) top view (blue plane) and side view (red plane). Detailed views are shown in Fig. 27. 5.2.5 Machining The goal of the engine design was to limit the amount of machining required because heat treatment increases the difficulty of Inconel 718 machining. The features that absolutely required machining were connection features, sealing grooves and alignment features. With surplus machining features on this material type, there is a lot of accrued cost when it comes to tooling. Tools used for cutting and drilling easily get worn down and need constant replacement, which adds time to machine setup. A wax will be employed to intentionally clog the internal channels prior to machining the critical features. By pooling melted wax into the internal sections of the nozzle and chamber, it will prevent any foreign object debris (FOD) from getting inside the regenerative channels. If the channels somehow get Fig. 26. Two planes for orifices inspection on the pintle Fig. 27. Top (top left), side (top right) view of pintle orifices, detail top clogged, it can have a detrimental impact on the engine during cold-flows and hot fire testing. Therefore, by utilizing wax, any object that attempts to penetrate the internal features will get stuck and picked out before melting out the wax. For the external surfaces of the engine that cannot be machined, bead blasting will be used to clean out any printing defects. This mostly includes down skin caused by higher laser speeds when printing close to 45o structures. It is possible to print the features, but the high laser speeds leave a very rough surface finish that needs to be cleaned up because it can injure the handlers of the component. Also, it can affect the performance of the engine based on the location of the down skin. Fig. 28 Full assembly of Balerion before Machining 6. Testing The purpose of the testing campaign is to characterize the design parameters of the hardware at the component level, which can provide information on selecting the propellant inlet conditions at the hot fire test. The campaign includes water flow tests on
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(IAC), Bremen, Germany, 1-5 October 2018. Copyright ©2018 by the International Astronautical Federation (IAF). All rights reserved. IAC-18-F1.2.3 Page 14 of 17 individual components and subsystems, proof and leak tests, cold-flow test and eventually the hot fire test. The objective and method of each test will be explained extensively in the following section. 6.1 Component Water Flow Testing Subsystem water-flow testing is essential as it provides information on the performance of designed parts. The results can be interpreted to validate the design parameters and analysis performed, as well as adjust the testing conditions accordingly in the following testing campaign. In Balerion’s testing campaign, USC LPL’s water flow test stand will be utilized to perform the water flow testing [7]. The components that require water flow tests include the pintle injector, both LOX and fuel sides separately and together, the regenerative cooling channels and the film cooling orifices. 6.1.1 Pintle LOX Side Water Flow Test The objective of this test is to verify if the current pintle tip geometry will provide a sufficient pressure drop at the designed operating condition and validate the pressure drop analysis result from engine sizing and ANSYS. The water flow test stand can measure the mass flow rate and the pintle orifice upstream pressure, which can be utilized to derive the flow coefficient, Cv, and the discharge coefficient, Cd, of the pintle tip geometry. Both non-dimensional values can serve as a tool to determine the pressure drop across the orifices at operating conditions, but the discharge coefficient can further provide information in the future redesign of the pintle tip geometry. The Cv value was found using the following equation: 𝐶𝑣 = 𝑄√ 𝑆𝐺 ∆𝑝 (7) where the Cv is the flow coefficient, Q is the volumetric flow rate in GPM, SG is the specific gravity of the fluid, and ∆𝑃 is the pressure drop across the testing article. For the determination of the Cd value of the pintle tip, equation (9) was used. Fig. 29. The testing article assembly Pintle + Center Annulus Multiple tests with different upstream pressures have been performed to mitigate measurement uncertainty and result computation error. The average Cv and Cd values were found to be 3.714 and 0.689, with a standard deviation of 0.007 and 0.001 respectively. By substituting the operating condition into the Cv equation, the pressure drop across the pintle tip at hot fire test is estimated to be approximately 98.7 psi, which is 26% of the chamber pressure, higher than the previous predicted value, 18%. Some potential reasons for the unmatched result might be the friction loss from the unpredictable surface roughness from the 3D printing and un-modelled turbulence that cause the additional pressure drop. However, this result is deemed acceptable since it should not affect the performance of the engine. Also, with the acquired Cd value, the lab will have a better understanding in sizing the pintle tip geometry to obtain a desirable pressure drop. 6.1.2 Engine Fuel Side Water Flow Test The goal of this test is to characterize the design parameters on the engine’s fuel side flow. Watched parameters include Cv and Cd values across the injector’s fuel ring outlet and the regenerative cooling channels, as well as the mass flow rate through the film cooling orifices. The Cv and Cd data at the fuel ring outlet can ensure the fuel side injector has a pressure drop at the designed operating condition. For the regenerative channels, the data can offer information on the fuel side propellant inlet condition during hot fire testing. The film cooling orifices mass flow rate is the most crucial data to obtain before the hot fire test. Initially, 12% of total fuel flow was figured to be the desired amunt to be used as the film cooling coolant. In the post-printing section, the orifices were too small to be inspected by the CT scan, thus no information was available on whether the orifices were clogged or had shrunk from printing. By knowing the actual mass flow rate coming of the film cooling orifices, the firing operation conditions can be re-evaluated. By performing water flow tests with the full Balerion assembly on the water flow test stand, the Cv and Cd values at the injector fuel outlet and the regenerative cooling channels can be calculated. The flow diverter, shown in Fig. 30, was designed to temporarily store the main injector fuel flow, while isolating the film cooling flow which can be measured by the water flow test stand. This setup allows the determination of the film cooling mass flow rate and the calculation of the injector’s main fuel flow rate. The acquired Cv and Cd values can be substituted with the operating condition to determine the actual injector fuel side pressure drop and the propellant inlet condition at the firing. The film cooling mass flow rate can also be
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(IAC), Bremen, Germany, 1-5 October 2018. Copyright ©2018 by the International Astronautical Federation (IAF). All rights reserved. IAC-18-F1.2.3 Page 15 of 17 used to evaluate whether the cooling system can achieve its designed thermal performance. Fig. 30. Flow diverter in full engine assembly 6.1.3 Cone Formation It is critical to visualize the spray pattern of the pintle injector to determine if the flow from the LOX side and fuel side collide and form a spray cone which indicates mixing between the propellants. Also, this will provide a data point to determine the relationship between the total momentum ratio (TMR) and the spray angle, since currently the spray angle is presumed to be 60 degree with the designed TMR value of 1 due to momentum conservation [2], when in actuality the spray angle needs to be tested and verified. The injector testing unit, a testing tool, shown below in Fig. 31, is designed to visualize the spray cone. This test article is designed to have a similar fuel outlet geometry as the engine to simulate the fuel flow while allowing the actual pintle injector to be mounted on to mimic the flow pattern. This test will be conducted on the water flow test stand with varying flow rates on both sides to achieve a TMR of 1 to investigate the spray angle of the cone pattern. This information will aid the lab in the future when designing pintle injectors as well as provide modelling inputs in ANSYS to more accurately simulate combustion in the chamber. Fig. 31. Injector testing unit CAD model 6.2 Proof & Leak Test To ensure the seals in between components’ interfaces are effective, proof and leak test are conducted to identify the potential leakage points which can be evaluated and mitigated before the firing. The major locations to check for leaks are the pintle to center annulus interface, the center annulus to chamber interface, the engine flange interface, and the sensor ports on the engine assembly. For the full engine proof and leak test, a testing tool is developed to seal the throat internally so that the pressurization of the chamber and regenerative cooling channels is possible. This testing tool, named throat plug, is a modification of an off-shelf expansion plug that utilizes the preload with center annulus in addition to the friction between the expanding rubber and nozzle inner wall to seal and retain itself against the pressure load of the fluid. The section view of the throat plug with the engine is shown below in Fig. 32. Fig. 32. Throat plug assembly on engine For the pintle and center annulus interface, in addition to the regular proof and leak tests, a separate cryogenic leak test needs to be conducted because in the firing sequence, the interface will be pre-chilled to cryogenic conditions. The purpose of this cryogenic test is to investigate under cryogenic temperatures, if the o- ring seal will still be effective and hold pressure. A blank pintle, see Fig. 33, was designed and manufactured to allow for pressurization of the assembly and test the sealing at the interface. Fig. 33. Blank pintle with center annulus assembly CAD model
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(IAC), Bremen, Germany, 1-5 October 2018. Copyright ©2018 by the International Astronautical Federation (IAF). All rights reserved. IAC-18-F1.2.3 Page 16 of 17 6.3 Cold-Flow Test The cold flow test of the engine should be conducted after results from the aforementioned tests are reviewed and accepted. The objective of the cold flow test is to characterize the test stand’s feed system with the engine integrated to supply the desired inlet propellant conditions at the hot fire test. For the test, water will be used on the fuel side since its property is similar to that of Jet-A, while for the LOX side liquid nitrogen will be used to simulate the cryogenic condition the engine will experience. In addition to characterizing the feed system, the cold flow test is useful in determining the chill-down procedure. It is desired to have the LOX side injector at correct thermal condition so that when LOX exits the injector it remains in the liquid state. The cold flow test, with the help of the thermocouple placed upstream of the injector, will help to determine if the proposed chill- down procedure needs to be modified. Furthermore, a timing test needs to be conducted to determine the time taken between sending the valve actuation command and when both the propellants reach the chamber. This will provide information on setting the ignition system because ideally the propellants can be ignited now when they exit the injector into the chamber. If not, there is a higher potential for the engine to over-pressurize at ignition, known as a hard start, due to the pooling of propellants in the engine, which can damage or even destruct the engine. Therefore, it is critical to perform the timing test to finalize the firing sequence while minimizing the time difference between when propellants enter the chamber and when ignition starts. 6.4 Hot Fire Test The main objective of the hot fire test is to verify the engine performance including thrust, cooling effectiveness, combustion efficiency and general reliability. A pre-fire inspection will be conducted and documented to compare with a post-fire inspection of the engine to investigate if the engine’s inner wall burns through during the firing. The temperature sensors located outside of the engine can help in constructing a thermal map across the engine’s wall to verify if the regenerative cooling channels and the film cooling orifices achieve the desired performance. The four pressure transducers located around the chamber allow the data acquisition system to collect pressure oscillation during the firing, thus provide information about whether the combustion is stable or not. Initially, short-duration firings (3 sec) will be conducted to determine if the ignition sequence is proper and to characterize the start transient condition of the engine. In addition, the engine shutdown procedure needs to be verified to ensure the safety of both the engine and the test stand. The shutdown sequence needs to be proved as reliable and repeatable under different firing conditions. Then, longer-duration firing (6-15 sec) can be conducted to verify the engine’s steady-state performance. 7. Conclusion Through the progress made on Balerion, additive manufacturing has proven to be very advantageous compared to conventional manufacturing methods. It has provided the capabilities to significantly reduce process time while allowing for increased complexity. However, with any new method, there will always be limitations. The main being the building volume availability in commercially available 3D metal printers. Additionally, a design meant for additive manufacturing must comply with constraints such as overhang and shrinkage that would not be considered in traditional manufacturing methods. Though additive manufacturing has progressed quite rapidly over the years, more advances still need to be made before it is capable of completely replacing conventional manufacturing. Critical features such as sealing surfaces, threads, etc. still rely on traditional machining methods as additive manufacturing cannot meet the required tolerance and surface finish needed for optimal system performance. As a student group, LPL has made substantial progress on the design and manufacturing of a fully 3D printed engine. The testing section (Section 6) lists the next major steps in the design verification of the Balerion engine. LPL has accumulated a lot of knowledge during the printing process and will continue to do so during the machining and testing phases. The goal of LPL is to build upon the lessons learned as to increase the skillset of its members in order to design better systems in the future. As mentioned, since the testing campaign has yet to be conducted, a future publication will be submitted to conclude the work on the Balerion Development Engine. Acknowledgements We would like to thank Department of Astronautical Engineering at the USC Viterbi School of Engineering for supporting LPL with infrastructure and safety oversight. Next, we would like to thank USC Centre for Advanced Manufacturing (CAM) for enabling us to use metal additive technology and the USC Machine Shop for providing expertise for technical drawings and machining our components. For CT scan images we would like to thank Aerospace Corporation. Extensive numerical simulations were possible thanks to sponsorship from Ansys and Nimbix.
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(IAC), Bremen, Germany, 1-5 October 2018. Copyright ©2018 by the International Astronautical Federation (IAF). All rights reserved. IAC-18-F1.2.3 Page 17 of 17 Finally, we would like to thank the current USC LPL team and alumni network for their willingness to push the boundaries of what a student-led group is capable of accomplishing. References [1] Stiegemeier, Benjamin, Michael Meyer, and Ray Taghavi. "A thermal stability and heat transfer investigation of five hydrocarbon fuels." 38th AIAA/ASME/SAE/ASEE Joint Propulsion Conference & Exhibit. 2002. [2] Huber, Marcia L., et al. "Preliminary surrogate mixture models for the thermophysical properties of rocket propellants RP-1 and RP-2." Energy & Fuels 23.6 (2009): 3083-3088. [3] Stechman, R. Carl, Joelee Oberstone, and J. C. Howell. "Design criteria for film cooling for small liquid-propellant rocket engines“ Journal of Spacecraft and Rockets 6.2 (1969): 97-102. [4] Dressler, G., & Bauer, J. (2000). TRW pintle engine heritage and performance characteristics. In 36th AIAA/ASME/SAE/ASEE Joint Propulsion Conference and Exhibit (p. 3871). [5] Habiballah, M., & Yang, V. (2004). Liquid rocket thrust chambers aspects of modeling, analysis, and design. Reston, Va.: American Institute of Aeronautics and Astronautics, Inc. [6] Huzel, D. K. (1992). Modern engineering for design of liquid-propellant rocket engines (Vol. 147). AIAA. [7] Moruzzi, M. Fessl, J., “Liquid Rocket Engine Component Water-Flow Test Stand,” Journal of Propulsion & Power (not yet published). [8] EOS M 290. (n.d.). Retrieved from https://www.eos.info/eos-m290 [9] Ellis, Adam, Ryan Brown, and Neil Hopkinson. "The effect of build orientation and surface modification on mechanical properties of high speed sintered parts." Surface Topography: Metrology and Properties 3.3 (2015): 034005.
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