Università degli Studi della Basilicata

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1 Università degli Studi della Basilicata SCUOLA DI INGEGNERIA CORSO DI LAUREA IN INGEGNERIA MECCANICA Tesi di Laurea in Macchine e Sistemi Energetici Titolo tesi Development of LRE Cooling System Module in a Concurrent Engineering Approach Relatore: Prof. Aldo Bonfiglioli Correlatori: Dott. Raffaele Votta Ing. Gianpaolo Elia Laureando: Sabato Massimo Matricola: ANNO ACCADEMICO 2013/14

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3 Table of Contents LIST OF FIGURES... II LIST OF TABLES... IV ACRONYMS, SYMBOLS AND ABBREVIATION... IV INTRODUCTION CHAPTER LIQUID ROCKET ENGINE OVERVIEW PROPULSION FUNDAMENTALS MAIN SYSTEMS OVERVIEW Feed System Thrust chamber Cooling System CHAPTER THE HYPROB PROGRAM & CONCURRENT DESIGN FACILITY THE HYPROB PROGRAM Introduction to National 2020 Vision Industrial heritage and program road map Propulsion Lines: LOx/LCH4 & Hybrid Design and measurement methodologies Experimental Facilities: CIRA & AVIO synergy CONCURRENT DESIGN FACILITY - CDF An innovative team working method History and status of CDF Applications, Benefits and key elements of CDF CHAPTER CIRA CONCURRENT DESIGN FACILITY CIRA CDF FOR SPACE PROPULSION ARCHITECTURE MODULE THRUST CHAMBER MODULE CHAPTER COOLING SYSTEM MODULE OF CIRA CDF OVERVIEW COOLING SYSTEM MODULE Heat flux analyses Cooling Channels geometry Coolant flow analysis CHAPTER RESULTS I

4 5.1 OVERVIEW VALIDATION OF THE COOLING SYSTEM DESIGN CYCLE CASE STUDY: 100 KN THRUST CLASS ENGINE Models of friction factor Cooling channels diameter effect Heat flux evaluation model CHAPTER CONCLUSIONS APPENDIX A SUPERCRITICAL FLUIDA S COOLANT IN LRE REFERENCES List of Figures FIGURE 1: GAS GENERATOR CYCLE, OPEN BLEED EXPANDER CYCLE & COMBUSTION TAP- OFF CYCLE FIGURE 2: STAGED-COMBUSTION CYCLE AND EXPANDER CYCLE FIGURE 3: PRESSURE-FED CYCLE FIGURE 4: THRUST CHAMBER SKETCH FIGURE 5: COAXIAL ELEMENT AND SHOWER HEAD FIGURE 6: UNLIKE DOUBLET AND UNLIKE TRIPLET FIGURE 7: LIKE-IMPINGING DOUBLET FIGURE 8: CONICAL NOZZLE FIGURE 9: PARABOLIC APPROXIMATION OF BELL NOZZLE CONTOUR FIGURE 10: ESA CDF IN SESSION FIGURE 11: LIQUID SPACE PROPULSION ENGINE - SCHEMATIC VIEW FIGURE 12: CIRA CONCURRENT DESIGN FACILITY FOR SPACE PROPULSION - SPECIALIST FIGURE 13: CIRA CONCURRENT DESIGN FACILITY FOR SPACE PROPULSION - DOMAINS FIGURE 14: CIRA CONCURRENT DESIGN FACILITY FOR SPACE PROPULSION - THE PROCESS FIGURE 15:TYPICAL BASIC CONFIGURATION OF A THRUST CHAMBER FIGURE 16: CONICAL NOZZLE CONTOUR FIGURE 17: BELL NOZZLE CONTOUR FIGURE 18: INITIAL AND FINAL PARABOLIC ANGLES VERSUS DESIRED NOZZLE EXPANSION RATIO FOR DIFFERENT PERCENT BELL LENGTHS OF AN EQUIVALENT 15 CONICAL NOZZLE FIGURE 19: CONVERGENT NOZZLE CONTOURS FOR STRAIGHT AND CUBIC SOLUTIONS FIGURE 20: AN EXAMPLE OF THE THRUST CHAMBER GEOMETRY EVALUATED BY THE ARCH MODULE FIGURE 21: REGENERATIVE COOLING ARCHITECTURE FIGURE 22: HEAT TRANSFER FOR SCHEMATIC REGENERATIVE COOLING FIGURE 23: VARIATION OF THERMAL CONDUCTIVITY WITH TEMPERATURE FOR TYPICAL METALLIC ELEMENTS AND ALLOY FIGURE 24: DETAIL VIEW COOLING CHANNEL GEOMETRY FIGURE 25: CROSS-SECTIONAL VIEW OF A REGENERATIVE COOLING THRUST CHAMBER SHOWING THE FLOWS DIRECTIONS FIGURE 26: TYPICAL CROSS-SECTIONAL SCALING OF A COOLING CHANNELS ALONG AXIAL DIRECTION FIGURE 27: MOODY DIAGRAM FIGURE 28: GEOMETRICAL PROFILE OF THRUST CHAMBER II

5 FIGURE 29: ARCHITECTURE CONCEPT FIGURE 30 COUNTER FLOW ARCHITECTURE FOR THE COOLING JACKET FIGURE 31: GEOMETRIC PROFILE OF THRUST CHAMBER FIGURE 32: COOLING SYSTEM CHANNEL AND BRAZING INTERFACE FIGURE 33: HEAT FLUXES GIVEN AS INPUT FIGURE 34: HYPROB-BREAD PRESSURE DISTRIBUTION VS CDF PRESSURE DISTRIBUTION FIGURE 35: HYPROB-BREAD TEMPERATURE DISTRIBUTION VS CDF TEMPERATURE DISTRIBUTION FIGURE 36: HYPROB-BREAD HEAT SPECIFIC DISTRIBUTION VS CDF HEAT SPECIFIC DISTRIBUTION FIGURE 37: HYPROB-BREAD THERMAL CONDUCTIVITY DISTRIBUTION VS CDF THERMAL CONDUCTIVITY DISTRIBUTION FIGURE 38: COOLING CHANNELS OF DEMONSTRATOR FIGURE 39: HEAT FLUX CFD ANALYSIS VS CDF FIGURE 40: TEMPERATURE CFD ANALYSIS VS CDF FIGURE 41: TEMPERATURE CFD ANALYSIS VS CDF FIGURE 42: SPECIFIC HEAT CFD ANALYSIS VS CDF FIGURE 43: THERMAL CONDUCTIVITY CFD ANALYSIS VS CDF FIGURE 44: GEOMETRICAL CONFIGURATION OF 100 KN CLASS THRUST CHAMBER FIGURE 45: HEAT FLUX DISTRIBUTION ALONG THE COOLING CHANNELS FIGURE 46: COOLANT DENSITY DISTRIBUTION ALONG THE COOLING CHANNELS FIGURE 47: COOLANT TEMPERATURE DISTRIBUTION ALONG THE COOLING CHANNELS FIGURE 48: COOLANT PRESSURE DROP DISTRIBUTION ALONG THE COOLING CHANNELS FIGURE 49: FRICTION FACTOR DISTRIBUTION ALONG THE COOLING CHANNELS FIGURE 50: SPECIFIC HEAT AT CONSTANT PRESSURE DISTRIBUTION ALONG THE COOLING CHANNELS FIGURE 51: COOLANT VELOCITY DISTRIBUTION ALONG THE COOLING CHANNELS FIGURE 52: REYNOLDS NUMBER DISTRIBUTION ALONG THE COOLING CHANNELS FIGURE 53: DYNAMIC VISCOSITY DISTRIBUTION ALONG THE COOLING CHANNELS FIGURE 54: THERMAL CONDUCTIVITY DISTRIBUTION ALONG THE COOLING CHANNELS FIGURE 55: COOLANT PRESSURE DISTRIBUTION ALONG THE CHANNELS FIGURE 56: DIAMETER EFFECT ON THE COOLANT VELOCITY ALONG THE COOLING CHANNELS FIGURE 57: DIAMETER EFFECT ON THE COOLANT PRESSURE DROP ALONG THE COOLING CHANNELS FIGURE 58: REYNOLDS NUMBER DISTRIBUTION ALONG THE COOLING CHANNELS FIGURE 59: DIAMETER EFFECT ON THE HEAT FLUX DISTRIBUTION FIGURE 60: DIAMETER EFFECT ON THE CONVECTIVE HEAT FLUX COEFFICIENT OF COOLANT DISTRIBUTION FIGURE 61: DIAMETER EFFECT ON THE GLOBAL COEFFICIENT OF HEAT TRANSFER DISTRIBUTION FIGURE 62: DIAMETER EFFECT ON THE TEMPERATURE DISTRIBUTION FIGURE 63: DIAMETER EFFECT ON THE COOLANT DENSITY DISTRIBUTION FIGURE 64: DIAMETER EFFECT ON THE COOLANT HEAT SPECIFIC DISTRIBUTION FIGURE 65: DIAMETER EFFECT ON THE COOLANT THERMAL CONDUCTIVITY DISTRIBUTION FIGURE 66: HEAT FLUX DISTRIBUTION ALONG THE COOLING CHANNELS FIGURE 67: COMPARISON BETWEEN HEAT FLUXES FIGURE 68: COMPARISON BETWEEN TEMPERATURE TRENDS FIGURE 69: COMPARISON BETWEEN DENSITY TRENDS FIGURE 70: COMPARISON BETWEEN VELOCITY TRENDS FIGURE 71: COMPARISON BETWEEN REYNOLDS NUMBER TRENDS FIGURE 72: COMPARISON BETWEEN SPECIFIC HEAT TRENDS FIGURE 73: COMPARISON BETWEEN THERMAL CONDUCTIVITY TRENDS FIGURE 74: COMPARISON BETWEEN PRESSURE TRENDS FIGURE 75: H2 AND CH4 COOLING CHANNEL OPERATIONAL CONDITION, ON A TYPICAL REDUCED PRESSURE-TEMPERATURE STATE DIAGRAM FIGURE 76 SPECIFIC HEAT AND THERMAL CONDUCTIVITY AS FUNCTION OF TENMPERATURE; P=6.0 MPA III

6 List of tables TABLE 1: ENGINE CYCLE ADVANTAGES AND DISADVANTAGES TABLE 2: NUMBER OF CHARACTERISTIC LENGTHS OF TYPICAL PROPELLANT COMBINATIONS TABLE 3: MAIN GEOMETRIC PARAMETERS OF HYPROB-DEMONSTRATOR TABLE 4: MAIN PERFORMANCE PARAMETERS OF HYPROB-DEMONSTRATOR TABLE 5: MAIN PERFORMANCE PARAMETERS Acronyms, symbols and abbreviation A A ch A t A 2 A w AR ARCH α ΔV CDF CFD COOL c CH 4 C p d d ch D t D c Nozzle section Channels section Throat section Exit area of nozzle Wetted area Aspect ratio Architecture system Nozzle divergence angle Delta Velocity Concurrent Design Facility Computational Fluid Dynamic Cooling system Characteristic Velocity Methane Heat Specific Diameter Diameter of cooling channels in the throat region Throat diameter Combustion chamber diameter IV

7 D H D e e ԑ FEED F f f s ξ g g 0 γ ρ h g h c H I sp INJE LRE L λ k m fu ch m fu m g L cham L conv L div M Hydraulic diameter Exit diameter Roughness Theoretical nozzle expansion area ratio Feed system Thrust Friction factor Safety factor Joint coefficient Acceleration of gravity Acceleration of gravity at sea level Specific heat ratio Density Convective Heat Flux Coefficient (gas) Convective Heat Flux Coefficient (coolant) Global coefficient of heat transfer Specific impulse Injection system Liquid Rocket Engine Characteristic Length Divergence angle correction factor for conical nozzle exit Thermal conductivity Mass flow rate (fuel) Mass flow rate (fuel) per unit channels Mass flow rate (gas) Chamber length Convergent nozzle length Divergent nozzle length Mach Number V

8 M x μ n N Nu O x O/F Pa p 2 p 3 p x p e p c Pr q r r e r t r x R R R Re R c R t R e R&T s t Local Mach Number Viscosity Number of the channels Newton Nusselt number Oxygen Mixture ratio Pascal Rocket gas pressure at nozzle exit Ambient or atmospheric pressure local gas pressure External pressure Chamber pressure Prandtl number Heat flux Local Recovery Factor Nozzle radius at the exit Nozzle radius at the throat Local nozzle radius Effective Recovery factor Gas constant per unit weight Universal gas constant Reynolds number Combustion chamber radius Throat radius Exit radius Research and Technology Distance between the cooling channels Chamber Wall Thickness VI

9 TCHA T x T y T aw T c (T c ) ns T co T wc T wg σ σ y v v x v v 2 V V c V conv Thrust chamber system Gas temperature at the section x Gas temperature at the section y Adiabatic wall temperature Chamber temperature Nozzle stagnation temperature Coolant Bulk temperature Coolant side wall temperature Gas side wall temperature Bartz correction factor for property variation across the boundary layer Yield stress Specific volume; Specific volume at section x Velocity Exit gas velocity Volume Combustion chamber volume Convergent nozzle volume V cc θ conv θ div θ n θ e Volume V c + V cc Converging nozzle angle Divergent half-cone angle Nozzle bell starting angle Nozzle lip exit angle VII

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11 INTRODUCTION This work of thesis has been carried out after a trainee period spent at the Propulsion department of the Italian Aerospace Research Center (CIRA), located in Capua, Italy. CIRA was created in 1984 to manage PRORA, the Italian Aerospace Research Program and uphold Italy s leadership in Aeronautics and Space. Among others, one of the most innovative research fields, in which CIRA is involved, is the aerospace propulsion. In particular, the activities aim at technological development for the modeling of rocket and ramjet/scramjet engines. The propulsion department is currently managing and working within the HYPROB program. As defined by the Italian Space Agency (ASI), this program will contribute to the implementation of national strategies for space propulsion. The strategic goal is to evolve and consolidate national technology and system development capabilities on rocket propulsion for future space applications. A detailed description of the program will be given in paragraph 2.1. One of the main product under development in the frame of HYPROB program is the CIRA Concurrent Design Facility (CDF) for Space Propulsion. The CDF exploits concurrent engineering methodology to perform effective, fast and cheap space mission preliminary studies. Concurrent Engineering Approach is the state-of-art methodology for the preliminary design phase of an Aero-Space Project (Phase 0/A). Equipped with a state-of-the-art network of computers, multimedia devices and software tools, the CDF allows team of experts to perform design studies during relatively short working sessions. The CDF design room has been designed and will be equipped with relevant hardware and software tools, with the aim of creating an effective communication and data interchange among team members. Concurrent Engineering Approach, along with CIRA CDF for Space Propulsion will be deeply described in paragraph 3.1. The present work is focused on the development of the Cooling System Module for a Liquid Rocket Engine. The cooling domain of the CDF will be equipped with this - 1 -

12 module and thus this work will allow the specialist to preliminary design and understand the feasible configurations of the cooling system. Thermofluidynamic behavior of the coolant along the cooling channels has been evaluated using engineering formulas and several approaches have been compared. In particular, a numerical investigation has been performed to determine the effect of the number of cooling channels on temperature, pressure drop and other thermofluidynamic properties of the coolant. Finally a validation of the performed work will be shown. This final result has been obtained comparing the developed module simulation with the results achieved by a 3 ton class LOx/CH4 LRE developed at CIRA in the framework of HYPROB Program

13 Chapter 1 LIQUID ROCKET ENGINE 1.1 Overview In this introductory chapter, the functioning of a LRE and its major components will be presented in detail. This description follows the key points proposed in literature. Rocket propulsion is a class of jet propulsion that produces thrust by ejecting stored matter, called the propellant. The energy from a high-pressure combustion reaction of propellant chemicals, usually a fuel and an oxidizing chemical, permits the heating of reaction product gases to very high temperatures (2200 to 3800 K). These gases are subsequently expanded in a nozzle and accelerated to high velocities (1800 to 4300 m/sec). Since these gas temperatures are about twice the melting point of steel, it is necessary to cool or insulate all the surfaces that are exposed to the hot gases. According to the physical state of the propellant, there are several different classes of chemical rocket propulsion devices. Liquid propellant rocket engines use liquid propellants that are fed under pressure from tanks into a thrust chamber. The liquid bipropellant consists of a liquid oxidizer (e.g., liquid oxygen) and a liquid fuel (hydrogen, kerosene, methane). A monopropellant is a single liquid that contains both oxidizing and fuel species; it decomposes into hot gas when properly catalyzed. Gas pressure feed systems are used mostly on low thrust, low total energy propulsion systems, such as those used for attitude control of flying vehicles, often with more than one thrust chamber per engine. Pump-fed liquid rocket systems are typically used in applications with larger amounts of propellants and higher thrusts, such as in space launch vehicles. In the thrust chamber the propellants react to form hot gases, which in turn are accelerated and ejected at a high velocity through a supersonic nozzle, thereby imparting momentum to the vehicle. A nozzle has a converging section, a constriction or throat, and a conical or bell-shaped diverging section. A liquid rocket propulsion system requires several precision valves and a complex feed - 3 -

14 mechanism which includes propellant pumps, turbines, or a propellant-pressurizing device, and a relatively intricate combustion or thrust chamber [1]. The present work is based on liquid bipropellant rocket engine. In particular, the attention has been focused on liquid oxygen as oxidizer and liquid methane as fuel. In order to better understand the functioning and the physics of rockets, next paragraph will deal with the basic equations that describe the performance parameters of a liquid rocket engine. 1.2 Propulsion Fundamentals The total impulse I t is the thrust force F, which can vary with time, integrated over the burning time t. I t is provided by Eq. (1.2 1). t I t = Fdt 0 (1.2 1) For constant thrust and negligible start and stop transients this reduces to Eq. (1.2 2). I t = Ft (1.2 2) The specific impulse I sp is the total impulse per unit weight of propellant. It is an important figure of merit of the performance of a rocket propulsion system. I sp can be expressed by Eq. (1.2 3). I sp = t Fdt 0 g 0 m dt (1.2 3) This equation will give a time-averaged specific impulse value for any rocket propulsion system, particularly where the thrust varies with time. During transient conditions, for instance during start or the thrust buildup period, the shutdown period, or during a change of flow or thrust levels, values of I sp can be obtained by integration or by determining average values for F and m for short time intervals. Considering constant thrust F and propellant flow m, and negligible short start or stop transients, this equation can be simplified as shown in the equation (1.2 4) reported below

15 I sp = F = F m g 0 w = I t m p g 0 = I t w (1.2 4) Where, m p is the total effective propellant mass, the product m p g0 is the total effective propellant weight, and ẁ is the weight flow rate. The concept of weight relates to the gravitational attraction at or near sea level, but in space or outer satellite orbits, "weight" signifies the mass multiplied by an arbitrary constant, namely g 0. In the Systeme International (SI) or metric system of units I sp can be expressed simply in seconds. However, the units of I sp do not represent a measure of elapsed time, but a thrust force per unit weight flow rate. In a rocket nozzle the actual exhaust velocity is not uniform over the entire exit cross-section and does not represent the entire thrust magnitude. The velocity profile is difficult to measure accurately. For convenience a uniform axial velocity c is assumed which allows a one-dimensional description of the problem. This effective exhaust velocity c is the average equivalent velocity at which propellant is ejected from the vehicle. It is defined by Eq. (1.2 5). c = I sp g 0 (1.2 5) It is given either in meters per second or feet per second. Since c and I sp differ only by an arbitrary constant, either one can be used as a measure of rocket performance. It is worth to note that the thrust is the force produced by a rocket propulsion system acting upon a vehicle. In a simplified way, it is the reaction experienced by its structure due to the ejection of matter at high velocity. It represents the same phenomenon that pushes a garden hose backwards or makes a gun recoil. The thrust and the mass flow are constant and the gas exit velocity is uniform and axial. In particular, this force is defined as: F = dm dt v 2 = m v 2 = w g 0 v 2 (1.2 6) This force represents the total propulsion force when the nozzle exit pressure equals the ambient pressure. Because of fixed nozzle geometry and changes in ambient - 5 -

16 pressure due to variations in altitude, there can be an imbalance of the external environment or atmospheric pressure p 3 and the local pressure p 2 of the hot gas jet at the exit plane of the nozzle. Thus, for a steadily operating rocket propulsion system moving through a homogeneous atmosphere, the total thrust is expressed by Eq. (1.2 7). F = m v 2 + (p 2 p 3 )A 2 (1.2 7) The first term is the momentum thrust represented by the product of the propellant mass flow rate and its exhaust velocity relative to the vehicle. The second term represents the pressure thrust consisting of the product of the cross-sectional area at the nozzle exit A 2 and the pressure difference evaluated at the same position. When the ambient atmosphere pressure is equal to the exhaust pressure, the pressure term is zero and the thrust is the same as in Eq. (1.2 6). In the vacuum of space p 3 = 0 and the thrust becomes: F = m v 2 + p 2 A 2 (1.2 8) The pressure condition in which the exhaust pressure is exactly matched to the surrounding fluid pressure (p 2 = p 3 ) is referred to the rocket nozzle with optimum expansion ratio. The effective exhaust velocity as defined by Eq. (1.2 5) applies to all rockets that thermodynamically expand hot gas in a nozzle and, indeed, to all mass-expulsion systems. From the previous equations its trivial to obtain that, for constant propellant mass flow, the exhaust velocity can be written as: c = v 2 + (p 2 p 3 ) A m 2 (1.2 9) Equation (1.2 10) shows that c can be determined from thrust and propellant flow measurements. When p 2 = p 3 the effective exhaust velocity c is equal to the average actual exhaust velocity of the propellant gases v 2. When p 2 p 3 then c v 2. The second term of the right-hand side of Eq. (1.2 9) is usually small in relation to v 2, thus the effective exhaust velocity is usually close in value to the actual exhaust velocity. The characteristic velocity has been used frequently in rocket propulsion literature. Its symbol c, is defined by Eq. (1.2 11). c* = p ca t m (1.2 11) - 6 -

17 The characteristic velocity c, is used in comparing the relative performance of different chemical rocket propulsion system designs and propellants; it is easily determined from measured data of m, p c, and A t. For ideal rocket, the hot gases behavior is described by some fundamental principles such as: Perfect gas law, defined by Eq. (1.2 12). p x vx = R T x (1.2 12) The principle of conservation of energy, defined as in Eq. (1.2 13). 1 2g 0 (v x 2 v y 2 ) = C p (T y T x ) (1.2 13) Principle of conservation of matter is expresses by the Eq. (1.2 14). m g = ρ x v x A x = constant (1.2 14) Finally, Eq. (1.2 15) that provides the Isentropic-Flow process. p x vx γ = constant (1.2 15) By an appropriate combination of these principles F, c* and other parameters can be written in the way reported hereinafter. Of course, the demonstrations is widely reported in literature[1] [5]. Area law can be expresses by Eq. (1.2 16) A y A x = M x M y { 1+ (γ 1) 2 M y 2 1+ (γ 1) 2 M x 2} (γ+1) (γ 1) (1.2 16) The exhaust velocity can be evaluated as : v 2 = 2γ R T γ 1 M c [1 ( p γ 1 2 ) γ ] (1.2 17) p 1 It can be seen that the exhaust velocity of a nozzle is a function of the pressure ratio p 2 p 1, the ratio of specific heats γ, and the absolute temperature at the nozzle inlet T c, as well as the gas constant R. Because the gas constant for any particular gas is inversely proportional to the molecular mass M, the exhaust velocity or the specific impulse are a function of the ratio of the absolute nozzle entrance temperature divided by the molecular mass. This ratio plays an important role in optimizing the - 7 -

18 mixture ratio in chemical rockets. In the latter equation the combustion chamber velocity has been considered negligible. Finally, the thrust can be expressed as: (γ+1) γ 1 F = A t p c 2γ2 ( 2 ) (γ 1) [1 ( p 2 ) γ ] + (p γ 1 γ+1 p 2 p 3 )A 2 (1.2 18) 1 The thrust, moreover, can be expressed in function of thrust coefficient C F defined as the thrust divided by the chamber pressure p c and the throat area A t. After some passages C F can be expresses by Eq. (1.2 19). (γ+1) C F = 2γ2 ( 2 ) (γ 1) [1 ( p γ 1 2 γ (p ) γ 1 γ+1 p ] + 2 p 3 )A 2 (1.2 19) 1 p c A t Therefore, the thrust becomes: F = C F A t p c (1.2 20) These equation provides an estimation of the performance of a liquid rocket engine. Of course, some of these equations have been used in the present work. The successive paragraphs deals with the main components of a LRE. 1.3 Main systems overview The overall architecture of a liquid rocket engine (LRE) is composed by several main systems, such as: Feed system: provide propellants to the injectors at the design pressure. It consists of: propellant tanks, pumps, turbines, valves and piping; Thrust chamber: is the key subassembly of a rocket engine. Here the liquid propellants are metered, injected, atomized, vaporized, mixed, and burned to form hot reaction gas products, which in turn are accelerated and ejected at high velocity[1]. A rocket thrust chamber is composed by injectors, combustion chamber, supersonic nozzle and mounting provisions. The injector has to introduce and meter the flow of liquid propellants to the combustion chamber, which provides an area for proper mixing of propellants and enough length to complete chemical combustion. The nozzle is - 8 -

19 responsible for the enthalpy conversion into kinetic energy and, thus, of the thrust generation[2]; Cooling system: this subsystem is mandatory due to the extremely high temperature reached by the thrust chamber walls. Regenerative cooling is the most commonly cooling technique used for LRE. It consists in let the coolant running through passages formed either by constructing the chamber liner from tubes or by milling channels in a solid liner[3]. This system is called regenerative because the coolant is the fuel itself. A deeper description is reported in the next sections of the paragraph Feed System In liquid bipropellant rocket engine systems, propellants are stored in one or more oxidizer tanks and one or more fuel tanks. A feed mechanism aims to move the propellants from tanks into the thrust chamber and raise propellants pressure. The energy for these functions comes either from a gas pressure feed system or turbo-pump feed system, thus a description of those two systems is described hereinafter. Finally, this paragraph will provide the main advantages and disadvantages of these feed system. Turbo-pump feed system and cycle engine: The propellants are pressurized by means of pumps, which in turn are driven by turbines. These turbines derive their power from the expansion of hot gases. Engines with turbo-pumps are preferred for booster and sustainer stages of space launch vehicles, long-range missiles, and in the past also for aircraft performance augmentation. Those systems are usually lighter than other types for these high thrust, long duration applications. The inert hardware mass of the rocket engine (without tanks) is essentially independent of duration. An engine cycle for turbo-pump fed engines describes the specific propellant flow paths through the major engine components, the method of providing the hot gas to one or more turbines, and the method of handling the turbine exhaust gases. There are open cycles and closed cycles. Open denotes that the working fluid exhausting from the turbine is discharged overboard, after having been expanded in a nozzle of its own, or - 9 -

20 discharged into the nozzle of the thrust chamber at a point in the expanding section far downstream of the nozzle throat. In closed cycles or topping cycles all the working fluid from the turbine is injected into the engine combustion chamber to make the most efficient use of its remaining energy. In closed cycles the turbine exhaust gas is expanded through the full pressure ratio of the main thrust chamber nozzle, thus giving a little more performance than the open cycles, where these exhaust gases expand only through a relatively small pressure ratio[1]. Now a discussion follows on several common open and closed engine cycles and their characteristics: Gas generator cycle. (see Figure 1) Open cycle. Pumps increase the propellant pressure before they are injected into the thrust chamber. The turbine that actuates the pumps is driven by a hot gas generator which combusts propellant tapped off from the main feed lines after the pumps. After it has passed the turbine, the gas is dumped into the atmosphere, sometimes through smaller nozzles to generate additional thrust, or alternatively injected back in the thrust chamber at the end of the nozzle. The portion of the fuel which does not go to the gas generator, passes first the nozzle where it is used for cooling before being injected in gaseous state into the thrust chamber; Bleed expander engine cycle. (see Figure 1) Open cycle. Pumps increase the propellant pressure before they are injected into the thrust chamber. The turbine that actuates the pumps is driven by hot gaseous fuel after it has passed as a liquid the nozzle where it is used for cooling. The gaseous fuel is dumped into the atmosphere after it has passed the turbine. The thrust chamber uses the gaseous fuel which is not send to the turbine. Combustion tap-off cycle. (see Figure 1) Open cycle. Pumps increase the propellant pressure before they are injected into the thrust chamber. The turbine that actuates the pumps is driven by hot combustion gas which is tapped of from the thrust chamber

21 Figure 1: Gas generator cycle, Open bleed expander cycle & Combustion tap- off cycle Staged combustion engine cycle. (see Figure 2) Closed cycle. Pumps increase the propellant pressure before they are injected into the thrust chamber. The turbine that actuates the pumps is driven by a warm gas generator which combusts oxidizer tapped off from the main oxidizer feed lines after the pump and gaseous fuel after the liquid fuel has passed the pump and the nozzle where it is used for cooling. In the warm gas generator the combustion is incomplete and the generated gas is injected in the thrust chamber where it combusts, in the ideal case, completely. Expander engine cycle. (see Figure 2) Closed cycle. Pumps increase the propellant pressure before they are injected into the thrust chamber. The turbine that actuates the pumps is driven by hot gaseous fuel after it has passed as a liquid the nozzle where it is used for cooling. After the gaseous fuel has passed the turbine it is injected into the thrust chamber[4]

22 Figure 2: Staged-combustion cycle and Expander cycle Gas pressure feed system: One of the simplest and most common means of pressurizing the propellants is to force them out of their respective tanks by displacing them with high-pressure gas. This gas is fed into the propellant tanks at a controlled pressure, thereby giving a controlled propellant discharge. Because of their relative simplicity, the rocket engines with pressurized feed systems can be very reliable. It consists of a high-pressure gas tank, a gas starting valve, a pressure regulator, propellant tanks, propellant valves, and feed lines. Additional components, such as filling and draining provisions, check valves, filters, flexible elastic bladders for separating the liquid from the pressurizing gas, and pressure sensors or gauges, are also often incorporated[1]. The functioning of a gas pressure feed system can be schematized as already done for the engine cycles. A short description on the gas pressure feed system is reported hereinafter: Pressure fed engine cycle. (see Figure 3) Closed cycle. No pumps are present, the oxidizer and fuel are injected directly in the thrust chamber. Because of the absence of pumps to increase the pressure after the tanks, the propellants have to be stored at high pressure [4]

23 Figure 3: Pressure-fed cycle Cycle advantages and disadvantages: In this paragraph the engine cycles will be compared to each other. Each cycle has its advantages and disadvantages in dry mass, wet mass, reliability, cost, specific impulse, etc. Table 1 shows such comparison

24 Cycle Advantages Disadvantages Pressure fed Gas generator Staged combustion Expander Bleed expander Simple reliable design No turbo-pump Fairly simple Wide thrust operating range High performance High chamber pressure and thrust capability Good performance Simple design with a low weight and wide thrust operating range No gas generator required No gas generator required Limited to low burn times and low thrust Limited throttling capabilities High pressure tanks Tank bladders can be required Turbine exhaust gas has low specific impulse and leads to effective loss in performance Gas generator required Very complex with lower reliability Advanced turbine and pumps required to cope with high pressures Pre-burner (gas generator) required Limited to low chamber pressures Limited to cryogenic fluids Table 1: Engine cycle advantages and disadvantages Limited to cryogenic fluids Pressure and thrust limited by fuel thermal properties The choice of a particular feed system depends on several parameters as propellants properties, material properties, mission time, chamber pressure and operating conditions. The present work has been developed considering a LRE with turbo-pump feed system and an regenerative expander cycle system Thrust chamber The thrust chamber is the heart of a propulsion system as it is the component which generates the thrust. As stated previously, this device is composed by an injection system, a combustion chamber, a supersonic nozzle and mounting provisions. The propellants are injected in the combustion chamber by injectors, here they react to form hot gases and develop large amounts of energy. The supersonic nozzle accelerates and ejects at high velocity the hot gases and it is responsible of the thrust

25 generation. The hot gases temperature can reach 3800 K and the chamber pressure is supposed to be as high as possible to increase the performances. Hence, this situation requires a careful design process. This section aims to describe in detail the components and the functions of the thrust chamber. In Figure 4 an example of a thrust chamber is reported. Injectors: Figure 4: Thrust chamber sketch The functions of the injector are similar to those of a carburetor of an internal combustion engine. The injector has to introduce and meter the flow of liquid propellants to the combustion chamber, cause the liquids to be broken up into small droplets (a process called atomization), and distribute and mix the propellants in such a manner that a correctly proportioned mixture of fuel and oxidizer will result, with uniform propellant mass flow and composition over the chamber cross section[1]. However, the injector, located directly over the high-pressure combustion, performs many other functions related to the combustion and cooling processes and is much more important to the function of the rocket engine than the carburetor is for an

26 automobile engine. No other component of a rocket engine has as great an impact upon engine performance as the injector. Injector design, like many engineering tasks, entails many compromises. The proper design starting point considers the particular application, engine size, propellant combination, and design priorities. Of course, the initial approach invokes complete optimization of all features: light weight, high performance, low cost, reliability, etc.; but that soon emphasizes priority for the main design parameters. One of most common problems relevant the injectors design is linked at combustion instability. All systems which release large amounts of energy have the potential for destructive oscillations, particularly if there is regenerative feedback (gain) between the combustion phenomena and the rate of energy release. This is particularly true of the combustion process, because temperature and pressure variations can directly impact the rates of vaporization and reaction. Stable operation can be achieved by either damping or detuning these processes. Hence, high performance can become secondary if the system can easily be triggered into a destructive instability, and many of the injector parameters that provide high performance appear to reduce the stability margin[5]. Now a discussion follows on the most common types of injection elements that are: non-impinging, unlike-impinging and like-impinging. Injection Elements: Nonimpinging elements Coaxial. The coaxial, or concentric, injection element usually has a slow-moving central stream of liquid oxidizer surrounded by a high-velocity concentric sheet of gaseous fuel. The liquid oxidizer is deliberately injected at low velocity, with the usual injection pressure-drop accomplished by an upstream metering orifice in each element, and diffused to a reduced velocity in the tubular LOx post. On the other hand, the fuel injection pressure is turned into high injection velocity in the annular gap around the LOx post. Mixing, atomization of the liquid, and mass distribution are provided by the shearing action of the high-velocity gaseous fuel on the surface of the liquid. The fuel surrounding the oxidizer tends to shield the combustion process, which enjoys a favorable combustor-wall heating environment, and also appears to benefit combustion stability

27 Showerhead. Directly axial, or near-axial, non-impinging streams of either liquid or gaseous propellants are generally referred to as "shower-heads." This type of element provides very little effective atomization or mixing, and is seldom used for primary injection. It is most frequently used for fuel-film-cooling streams at the chamber wall. There are other types of injection elements, such as Fan formers and Slots and sheets, but they have seldom been successful[5]. The injection elements which have just been described, are represented in Figure 5. Figure 5: Coaxial element and Shower head Injection Elements: Unlike-impinging elements Unlike doublets. A straightforward way of mixing two different fluid streams directs one against the other; this in essence describes the basic unlike-impinging doublet. The impact produces a fan-shaped spray made up of a mixture of the two impinging fluids. With no combustion or other chemical reactions, the combined streams form a largely two-dimensional spray in a plane basically at right angles to the plane which includes the centerlines of the impinging streams. The width of the spray fan largely reflects the included impingement angle of the two streams, the thickness to the stream diameters, and the turbulence level. Mixing in the spray fan is not perfectly distributed, being adversely affected by any momentum and/or stream-diameter mismatch of the impinging fluids. Stream misimpingement, resulting from the fact that the stream centerlines rarely intersect at the theoretical impingement point, distorts the shape of the spray fan and produces mixing imperfections. Other effects can be arise when combustion processes are superimposed upon impinging-stream hydrodynamics

28 Unlike triplets. A mismatch in stream size and momentum between the oxidizer and the fuel in unlike doublet elements will force the spray away from the desired axial direction and distort the fan, resulting in poorer mixing. This problem may be avoided by use of a symmetrical, unlike-injection element consisting of an axial central stream of one propellant and two symmetrically-impinging outer streams of the other propellant. This unlike triplet may have either two fuel streams impinging on a central oxidizer stream (fuel-oxidizer-fuel) or the reverse (oxidizer-fueloxidizer). In most propellant combinations, the total oxidizer flow area will be the greater, so the O-F-O system provides a closer match of stream sizes and consequently better mixing. Unlike-triplet injectors have demonstrated high levels of mixing and resultant combustion efficiency, but they also tend to be sensitive to stability problems[5]. An example of Unlike doublets and Unlike triplets are represented in Figure 6. Figure 6: Unlike doublet and Unlike triplet Injection Elements: Like-impinging elements Like doublets. Like-impinging elements impinge the injected streams (liquid or gas) directly on other streams of the same propellant. The most common of these, a doublet configuration, has two like-fluid streams angled together to an impact point, producing in a fan-shaped spray of droplets similar to that of an unlike doublet. There is no mixing within this fan, since only one reactant is present in each. Energy dissipated by the impingement atomizes the liquids. Like-impinging elements are frequently used for liquid/liquid propellant systems in which reaction or heat transfer between unlike-impinging streams is undesirable. The like-impinging doublet avoids most of the reactive-stream de-mixing of unlike-impinging designs and better maintains combustion stability than unlike patterns[5]

29 In addition, a triplet configuration have been developed in which three streams of the same propellants can be directed to a common impingement point. An example of a like doublets is reported in Figure 7. Combustion chamber: Figure 7: Like-impinging doublet A liquid-rocket combustion chamber converts propellants into high-temperature, high-pressure gas through combustion, which releases the chemical energy of the propellant, resulting in an increase in internal energy of the gas. The liquid propellants are injected at the injection plane with a small axial velocity which is assumed to be zero in gas-flow calculations. The combustion process proceeds throughout the length of the chamber and is expected to be completed at the nozzle entrance. Heat released between injection plane and nozzle inlet increases the specific volume of the gas. To satisfy the conditions of constant mass flow, the gas must be accelerated toward the nozzle inlet with some drop of pressure. The combustion temperature is much higher than the melting points of most chamber wall materials, therefore it is necessary either to cool these walls or to stop rocket operation before the critical wall areas become too hot. If the heat transfer is too high and thus the wall temperatures become locally too high, the thrust chamber will fail. Nowadays the preferred solution is composed by a cylindrical chamber with a flat injector and a converging-diverging nozzle. The chamber volume is defined as the volume up to the nozzle throat section and it includes the cylindrical chamber and the converging cone frustum of the nozzle. The volume and shape are selected after evaluating some constraints: The volume has to be large enough for adequate mixing, evaporation, and complete combustion of propellants. Chamber volumes vary for different propellants with the time delay necessary to vaporize and activate the

30 propellants and with the speed of reaction of the propellant combination. When the chamber volume is too small, combustion is incomplete and the performance is poor. With higher chamber pressures or with highly reactive propellants, and with injectors that give improved mixing, a smaller chamber volume is usually permissible. The chamber diameter and volume can influence the cooling requirements. If the chamber volume and the chamber diameter are large, the heat transfer rates to the walls will be reduced, the area exposed to heat will be large, and the walls are somewhat thicker. Conversely, if the volume and cross section are small, the inner wall surface area and the inert mass will be smaller, but the chamber gas velocities and the heat transfer rates will be increased. There is an optimum chamber volume and diameter where the total heat absorbed by the walls will be a minimum. This is important when the available cooling capacity of the coolant is limited (for example oxygen-hydrocarbon at high mixture ratios) or if the maximum permissive coolant temperature has to be limited (for safety reasons with hydrazine cooling). The total heat transfer can also be further reduced by going to a rich mixture ratio or by adding film cooling (a technique discussed below). All inert components should have minimum mass. The thrust chamber mass is a function of the chamber dimensions, chamber pressure, and nozzle area ratio, and the cooling method. Manufacturing considerations favor a simple chamber geometry, such as a cylinder with a double cone bow-tie-shaped nozzle, low cost materials, and simple fabrication processes. In some applications the length of the chamber and the nozzle relate directly to the overall length of the vehicle. A large-diameter but short chamber can allow a somewhat shorter vehicle with a lower structural inert vehicle mass. The gas pressure drop for accelerating the combustion products within the chamber should be a minimum; any pressure reduction at the nozzle inlet reduces the exhaust velocity and the performance of the vehicle. These losses become appreciable when the chamber area is less than three times the throat area

31 For the same thrust, the combustion volume and the nozzle throat area become smaller as the operating chamber pressure is increased. This means that the chamber length and the nozzle length (for the same area ratio) also decrease with increasing chamber pressure. The performance also goes up with chamber pressure[1]. The preceding chamber considerations conflict with each other. Depending on the application, a compromise solution that will satisfy the majority of these considerations is therefore usually selected and verified by experimental data. Nozzle: This paragraph will describe the functioning of nozzle and the hot gases behavior. As already told, the combustion products are discharged through a converging-diverging nozzle to achieve high gas velocities and thrust. This phenomena will be described in condition of ideal rocket. Such hypothesis allows to express the basic thermodynamic principles with simple mathematical relationships. Besides, the flow in the nozzle will be considered quasi-one-dimensional. Gas flow through rocket nozzles The prime function of a rocket nozzle is to convert efficiently the enthalpy of the combustion gases into kinetic energy and thus create high exhaust velocity of the gas. The nozzle is the most efficient device for accelerating gases to supersonic velocities. Rocket nozzles are conventionally of the converging-diverging De Laval type, with the cross-sectional area decreasing to a minimum at the throat and then increasing to the exit area. The flow velocity through a nozzle increases to sonic velocity at the throat and then increases further supersonically in the diverging section. In practice, for one-dimensional isentropic expansion, it is assumed that the gas flow through the nozzle will be an isentropic expansion, and that both the total temperature and the total pressure will remain constant throughout the nozzle. The static pressure at a nozzle throat with sonic flow, where the maximum weight flow per unit area occurs, is defined as critical pressure. The velocity of sound is equal to the velocity of propagation of a pressure wave within a medium. It is therefore impossible for a pressure disturbance downstream of the nozzle throat to influence

32 the flow at the throat or upstream of the throat, provided that this disturbance will not create a higher throat pressure than the critical pressure. It is one of the characteristic features of an attached diverging or De Laval nozzle, however, that sonic velocity in the nozzle throat is maintained even if the back pressure (ambient pressure) at the nozzle exit is greater than the pressure required at the throat for sonic velocity. As a result, a pressure adjustment (recovery) must take place between the throat and the nozzle exit (ambient pressure). This adjustment may take place through subsonic deceleration (isentropic) or by way of non-isentropic discontinuities called shock waves, or a combination of both. In short, pressures lower than ambient may be present in a supersonic nozzle. The higher ambient pressure cannot advance upstream within the nozzle, since the gases are flowing with supersonic velocity. An exception to this is in the region of the flow along the nozzle walls, where, due to friction, a boundary layer of slow-moving gases may exist. In this subsonic boundary layer, ambient pressure may advance for a distance, forcing the low-pressure center jet away from the walls. It might be expected that the point of separation will be at the point of optimum expansion, but separation usually occurs further down-stream. In fact, it rarely occurs at all in conventional rocket nozzles within the designed region of operation, unless an extreme case of overexpansion exists or unless excessive nozzle divergence angles are chosen[5]. Nozzle configuration A number of different proven nozzle configurations are available nowadays. The principal difference in the different nozzle configurations is found in the diverging supersonic-flow section. The wall surface throughout the nozzle should be smooth and shiny to minimize friction, radiation absorption, and convective heat transfer due to surface roughness. Gaps, holes, sharp edges, or protrusions must be avoided. The most common nozzle configurations are conical nozzle and bell-shaped nozzle. Conical nozzle. In early rocket-engine applications, the conical nozzle, which proved to be satisfactory in most respects, was used almost exclusively. A conical nozzle allows ease of manufacture and flexibility in converting an existing design to higher or lower expansion area ratio without major redesign. Since certain performance losses occur in a conical nozzle as a result of the non-axial component of the exhaust

33 gas velocity, a correction factor, is applied in the calculation of the exit-gas momentum[5]. This factor (thrust efficiency) is the ratio between the exit-gas momentum of the conical nozzle and that of an ideal nozzle with uniform, parallel, axial gas-flow. The value of this parameter can be expressed by the following equation ( ): λ = 1 2 (1 + cos α) ( ) The configuration of a typical conical nozzle is shown in Figure 9:. Figure 8: Conical nozzle Bell nozzle. To gain higher performance and shorter length, engineers developed the bell-shaped nozzle. It employs a fast-expansion (radial-flow) section in the initial divergent region, which leads to a uniform, axially directed flow at the nozzle exit. The wall contour is changed gradually enough to prevent oblique shocks. The expansion in the supersonic bell nozzle is more efficient than in a simple straight cone of similar area ratio and length, because the wall contour is designed to minimize losses[5]. One convenient way of designing a near-optimum-thrust bell nozzle contour uses the parabolic approximation procedures. The design configuration of a parabolic approximation bell nozzle is shown in Figure 9 shows the contour of a bell nozzle

34 Figure 9: Parabolic approximation of bell nozzle contour Cooling System All rocket engines show a common problem, high energy released by combusted gases. This problem results in high combustion temperatures (2200 to 3600 K), high heat transfer rates (0.8 to 160 MW/m 2 ) in thrust chamber and requires special cooling techniques for the engine. Cooling techniques developed to cope with this problem, either singly or in combination, include regenerative cooling, radiation cooling, film or transpiration cooling, ablation, arid inert or endothermic heat sinks. To choose the proper cooling technique mission requirements, environmental requirements and operational requirements should be considered. Regenerative cooling Regenerative cooling is performed building cooling jackets around the thrust chamber and circulating one of the liquid propellants, usually the fuel, through them before the fuel is fed to the injector plate[1]. Regenerative cooling is one of the most widely applied cooling techniques in liquid propellant rocket engines. It has been effectively applied with high chamber pressure systems and for long durations with a wide heat flux range, form 0.8 to 160 MW/m 2. Besides, this cooling technique is used primarily with bipropellant chambers and medium/large thrust. The structure is relatively light, however, regenerative cooling has also some disadvantages that include limited throttling with most coolants, reduced reliability with some coolants, high pressure drops required at high-heat-flux levels, and thrust levels, mixture ratios, or nozzle area ratios possibly limited by maximum allowable coolant-temperature

35 It is possible to think about regenerative cooling of a liquid propellant rocket engine as a balance between the energy rejected by the combusted gases and the heat energy absorbed by the coolant. The energy absorbed by the coolant is not wasted but it augments the initial energy content of the propellant prior to injection, slightly increasing the exhaust velocity (0.1 up to 1.5%). Therefore thermal energy is recovered in the system. However by this process the overall engine performance gain is less than 1% [3]. In particular, the bulk temperature of the coolant increases from the point of entry until it leaves the cooling passages, as a function of the heat absorbed and of the coolant flowrate. To maintain the chamber walls at temperatures below those at which failure might occur because of melting or stress, a proper balance of these parameters becomes of major importance for the design of a regeneratively cooled thrust chambers. For metals commonly used in thrust-chamber walls, such as stainless steel, nickel, NARLoy-Z, and nickel-base super-alloys, the limiting hot-gas-side wall temperature ranges from 700 to 1300 K. The resultant differences between combustion-gas temperature and wall temperature range from 1600 to 3600 K. Sometimes, regenerative cooling, with attendant pressure losses requiring additional turbopump power or higher gas-pressurization levels, imposes an overall performance penalty. Design of a regeneratively cooled thrust chamber involves consideration of gas-side heat flux, wall structural requirements, coolant-side heat transfer, and the effects of temperature increases on coolant properties[5]. Dump cooling Dump cooling. With this principle, a small percentage of the propellant, such as the hydrogen in a LO2/LH2 engine, is fed through passages in the thrust chamber wall for cooling and is subsequently dumped overboard through openings at the rear end of the nozzle skirt. Because of inherent problems, such as performance losses, this method has only limited application. Film cooling Here, exposed chamber-wall surfaces are protected from excessive heat by a thin film of coolant or propellant introduced through orifices around the injector

36 periphery or through manifolded orifices in the chamber wall near the injector and sometimes in several more planes toward the throat. The method has been used, particularly for high heat fluxes, either alone or in combination with regenerative cooling. Traspiration Cooling Transpiration cooling introduces a coolant (either gaseous or liquid propellant) through porous chamber walls at a rate sufficient to maintain the desired temperature of the combustion-gas-side chamber wall. This method is essentially a special type of film cooling. Ablative cooling In this process, combustion-gas-side wall material is sacrificed by melting, vaporization, and chemical changes to dissipate heat. As a result, relatively cool gases flow over the wall surface, thus lowering the boundary-layer temperature and assisting the cooling process. In addition, the ablative material is usually a good thermal insulator, keeping to a minimum the heat transmitted to the outer structure. Ablative cooling has been used in numerous designs, initially mainly for solidpropellant systems, but later, equally successfully, for short-duration and/or low- p c liquid systems. Radiation Cooling With this method, heat is radiated away from the surface of the outer thrust-chamber wall. It has been successfully applied to very small, high-temperature-material combustion chambers and to low-heat-flux regions, such as nozzle extensions[5]. This work of thesis is focused on the development of a regenerative cooling system module, thus chapter 4 will explain in detail the mathematic model describing the regenerative cooling and the thermo-fluidynamics properties of the coolant

37 Chapter 2 THE HYPROB PROGRAM & CONCURRENT DESIGN FACILITY 2.1 THE HYPROB PROGRAM One of the most important activity in which CIRA is involved is the HYPROB Program. CIRA Concurrent Design Facility for Space Propulsion is one of the main project of this program and this paragraph will provide a broad description of the activities managed by the propulsion department. The Italian program HYPROB, kicked-off in 2010, is carried out by CIRA under contract by the Italian Ministry of University and Research (MIUR), as contribution to the National Aerospace Research Program (PRORA), in coherence with the long-term vision of the Italian Space Agency on Space Propulsion and the needs of industrial national stakeholders. The program relies upon the national heritage resulting from other development programs, supported by the Italian Space Agency (ASI) at both national (LYRA) and European (FLPP) level, mainly focused on the evolution of launchers, and represents a R&T effort to contribute to further develop space propulsion assets at national level Introduction to National 2020 Vision Propulsion systems based on hydrocarbons, either liquid or hybrid, represent nowadays a major technology challenge for future launchers and space transportation systems, to be pursued through R&T demonstration programs addressing enabling

38 technologies. Methane is one of the most interesting solutions as propellant for liquid rocket engines, in combination with Oxygen, due to good performances achievable in terms of specific impulse (I sp ~ 380 s) combined with operation advantages, such as storability, low toxicity, availability and production cost, as compared to hydrogen. Additional features of methane regard its good cooling capability and well known material compatibility, that makes it ideal for regenerative thrust chambers. In a long term perspective, such a propulsion technology may encompass a wide range of propulsion systems, from launcher main stages up to small thrusters, but present envisaged applications regard mostly: upper stages of small launchers primary propulsion systems for interplanetary missions, such as ascent ad landing modules Hybrid technology also is of great interest for space propulsion, combining the best features of both solid, namely storability, and liquid option, namely performances. However, although the potential of such a technology has been proven, the level of maturity toward real applications has not been completely achieved yet. Possible architectures and features of future Space Transportation Systems are strongly influenced by technologies that are already available and others that require further or completely new developments. Furthermore, the actual scenario of European launcher family and of the worldwide sector, dealing with the threats of a highly competitive market, daily faces the challenge of increasing performances and reliability, in parallel with cost reduction. Propulsion disciplines constitute an asset of such space technology, especially in view of developing new skills leading to define possible evolutions and future generation launch vehicles and space transportation systems. Aiming to support and promote the consolidation and the evolution of competences in the field by the national scientific and industrial community, an integrated national program of research and development activities has been structured taking the maximum results from Ministry of Research and University initiatives and from ASI on going and future programs, then preparing for the future technical challenges. As far as chemical propulsion is concerned, the background gained by the Italian community is strongly based on solid rocket motors, that have mainly contributed to the success of Vega qualification flight in February 2012, which consolidation is one of the key elements of the future national

39 vision. In the Horizon 2020, strategic importance is given to the R&D (Research and Development) programs in liquid and innovative propulsion. Several programs have been started by years, mainly on LOx- Methane propulsion, with the development of an engine demonstrator for the upper stage of Vega evolutions, and the setup of a dedicated test facility. This activity is integrated by the HYPROB program, resulting in the acquisition of base research competences and engineering design skills up to the fully national development of the entire combustion chamber. Furthermore, activities of research and experimental demonstration on Hybrid propulsion will be pursued, in order to take the better results by the integration of competences in liquid and solid rocket design, leading to promising alternative solutions. The synergy among industry, research centers and university competences skills and infrastructures, is a key element of such vision, as well as it is the international cooperation with other space agencies, as actually are Roscosmos and JAXA Industrial heritage and program road map The main heritage at industrial level on methane-based propulsion relies on AVIO Group. Similarly to other worldwide industrial leaders in aerospace rocket design and fabrication, in recent years AVIO Group has been carrying out R&D activities in this field through either self-sponsored internal programs and projects sponsored by both the Italian (ASI) and the European (ESA) space agency. More specifically, investigations have regarded different aspects related to chemical rocket engines: LOx/CH 4 - combustion phenomena, through small-scale engines combustion chambers and torch, tested in FAST_2 facility; Inducer super-cavitation phenomena, through a test article design based on previous experience done by AVIO Group in the Ariane 5 Program (Vinci, Vulcain, etc); Characterization of hydrostatic bearings in a cryogenic environment (LN2); Verification of the coupling between hot, high pressure gas from a Test Burner working in a fuel rich environment and turbine stator vanes sample, assessing the risk of sooting

40 AVIO Group is also deeply involved in the national program LYRA, funded by the Italian Space Agency with the goal of developing technologies for future cryogenic upper stage propulsion. In this regard, a preliminary configuration, derived from VEGA launch vehicle, and using a new LOx-LNG upper stage has been defined aimed at improving the payload capabilities. A 100 kn demonstrator, representative of an expander cycle engine, named LM10-MIRA, has been developed. The design has been derived by the Russian KBKhA RD-0146 engine, combined with fuel turbo-pump and injection head design developed by AVIO. The system has been tested in Russian facilities with the aim of gaining additional experience on Oxygen-Methane propulsion technology. AVIO is also carrying out the THESEUS project, again supported by the Italian Space Agency, aimed at investigating thrusters evolution for space exploration. The main focus is on hybrid technologies and ablative cooling chambers. In the above described national framework, the HYPROB Program strategic objectives and the overall development plan have been set in a preliminary step, based on interactions with the institutional, industrial and scientific stakeholders. This step was completed in early 2011 with a Concept of Operation review. The main outcome of this step was to maintain the focus on both liquid oxygen-methane (LOx /LCH 4 ) and hybrid technology, in order to harmonise and consolidate the national heritage from previous R&D activities. In this respect, the program Road Map pursues a mid-term goal, in the time frame , related to the assessment of system capabilities and technologies at demonstration level, based on a technology-push approach, and a longer term goal, to be pursued in the time frame and beyond, where those technologies will be devoted a specific space application, based on a system-driven approach. In the mid-term perspective, for both liquid and hybrid developments, the focus is pointed at: development of technology demonstrators, including intermediate breadboards; development of R&D activities in relevant technology areas; improvement of test capabilities

41 At system level, the mid-term objective is to design, manufacture and test, in a relevant facility, technology demonstrators of suitable class of thrust, with the main scope of validating critical design and technology features and then to assess technology readiness level of potential solutions for future engines. In the framework of R&D activities, the focus is put on enabling technologies, such as combustion modeling, thermo-mechanical modeling, materials and manufacturing processes at both system and components level. Specifically in the methodology field, the main scope is to enhance the capabilities of simulating the complex combustion and thermo-mechanical processes, characteristic of both liquid and hybrid propulsion, as a fundamental step to improve the design processes for future applications. The models will be validated through extensive testing activities at small-scale level in either newly designed or up-graded test benches Propulsion Lines: LOx /LCH 4 & Hybrid The System line devoted to the LOx/LCH 4 technology aims at designing, manufacturing and testing a LRE ground demonstrator, representative of a 30 kn of thrust in flight conditions (vacuum). The architecture considered for the demonstrator, in line with the project key level requirements, is a regenerative cooled thrust chamber for ground testing. Regenerative cooling is one of the most widely applied cooling techniques used in liquid propellant rocket engines. As stated in the introductory chapter, it has been effective in applications with high chamber pressure and for long durations with a heat flux ranging from 1.6 to 160 MW/m 2. In particular, in expander engines, regenerative cooling enthalpy gain is used to move turbines for pressurizing pumps. The study logic implemented in that project has been based on the following drivers: Exploit existing know-how and design solutions for critical items; Design suitable intermediate breadboards to address the most critical design solutions, such as injection and cooling

42 This approach has been defined in order to proceed step by step, from the understanding of the basic physical processes, i.e. combustion and heat transfer, to the validation of design and analysis methodologies. The studies carried out in the program will benefit the collaboration between ASI and JAXA (Japan Aerospace Exploration Agency) agencies on methane technology. The System line devoted to the Hybrid technology has again the objective of developing a demonstrator of similar thrust class. The main interest is on the combination of paraffin with either oxygen or nitrogen-based oxidants. The selection of paraffin as solid propellant has been made due to the complementarities with other national developments where the HTPB (Hydroxyl-terminated polybutadiene) polymer has been considered. The activity is being carried out in collaboration with other national research institutions, in order to benefit of a solid scientific and technology background for demonstration purposes. The development plan has been set out in 2011 and approved after a Concept of Operations review. A first slice of the project is devoted to the development of R&D activities on both the oxidant and the propellant sides. This has yield to the selection of the technologies to be integrated into the demonstrator, based on suitable tests at sub-scale level Design and measurement methodologies As clarified in previous sections, the final goal is to improve the capabilities in the design of future rocket engines; to this aim, all the numerical tools developed or tested within the project will be properly interfaced one each other and will be used by following a concurrent design approach. The set-up of a small Concurrent Design Facility, in which the experts of different disciplines will be allowed to meet and work together to optimize the design phases, is also foreseen. The capability to perform a detailed analysis of the processes occurring inside the combustion chamber will be one of the fundamental aspects that will be taken into account; on one side an advanced CFD code, named SPARK (Solver for Propulsion Applications including Real Gas Kinetics), is being developed and validated, in order to be able to numerically simulate the fluid dynamics inside the combustion chamber. On the other hand, the code will be validated in representative conditions of a LRE

43 environment. The experimental data will be also made available from the testing activities and thus used for code validation. State of the art models have been implemented in the CFD code, in order to be able to take into account the most important relevant phenomena, including high pressure real gas effects, turbulent combustion, and sprays. A Large Eddy Simulation (LES) model will be implemented as well. SPARK code will be applicable also for regenerative cooling system simulation, that is one of the main goals of the system activities within the program. As far as experimental activities are concerned, a preliminary design phase has been carried out with the goal to identify the main diagnostic systems to be used in the test bench, in order to obtain reliable and accurate data useful for code validation and for a deep comprehension of the physical phenomena occurring in the combustion chamber. Several optical techniques have been taken into account: High Speed Camera (HSC) imaging, High-Speed Shadowgraphy/Schlieren, Time-Resolved High Resolution Optical Emission Spectroscopy (OES), Planar Laser Induced Fluorescence (PLIF), Particle Image Velocimetry (PIV), Coherent Anti-Stokes Raman Spectroscopy (CARS). Obviously, in order to use optical diagnostics, the designed combustion chambers will be provided with optical accesses. The attention has being initially focused on the first four techniques, for which a significant experience was already held by CIRA for several previous applications Experimental Facilities: CIRA & AVIO synergy In the frame of HYPROB program, the realization of a combustion laboratory at CIRA is foreseen, for assembly and integration of breadboards and basic testing of combustion. This area will be useful to perform experimental research activities on combustion chambers and several Fuels/Oxidizers combinations. Furthermore it will be aimed at developing advanced diagnostic techniques and methodologies to investigate specific aspect related to combustion and support the activities carried out in the program itself. The test bench shall allow testing test articles representative of small combustion chambers, provided with a limited number of injectors (up to three), and able to withstand combustion chamber pressures up to 7 MPa, both in

44 subcritical and supercritical conditions, for a maximum run-time of 30 seconds. Oxygen and Methane, both gaseous and liquid, are the selected propellants to be used. However, since the test bench has been designed to be quite flexible, other fuels could be used, when needed. To improve facility productivity, two test stands, using the same fluids storage tanks and piping, will be built, each one being able to be used with a test setup different than the other one. The test bench shall be widely used in the frame of basic R&D activities. Among them, the development of advanced diagnostic techniques (such as PLIF, High Speed Cameras, High speed Shadowgraphy/Schlieren, High speed Optical Emission Spectroscopy) is envisaged to be performed using this small facility. With this aim, the facility layout is designed in such a way to allow both the proper setup of the instrumentation, optics and lasers, and a comfortable and easy use of the same ones. A diagnostic laboratory, hosting the laser instrumentation, shall be located just close to the test fire area. Furthermore, according to the synergic approach adopted in the program, FAST_2 facility running at Colleferro (Rome) within the AVIO plant, has been selected as the main experimental facility to carry out tests on demonstrators and associated combustion breadboards. The facility has been developed in the FAST_2 program, funded by ASI some years ago in support to space transportation technologies and is presently co-owned by ASI and AVIO Group. The main characteristics of the actual configuration of the facility are hereinafter reported: LOx feeding line up to 10 kg/s at 200 bar (max pressure tank); GCH 4 feeding line up to 2 kg/s at 200 bar (max pressure tank); Cooling water feeding line up to 20 kg/s at 140 bar; Test cell for combustor testing; Command and control capability provided by two redundant units; Data acquisition unit with 82 channels. The facility has already been extensively used in recent programs to test thrust chambers representative of 30 kn thrust (vacuum) class. According to HYPROB

45 program requirements[6], a severe update of the facility will be carried out to enable testing capability with liquid methane. 2.2 CONCURRENT DESIGN FACILITY - CDF In recent years, the role of the CDF and the methodology adopted by this innovative designing strategy has become increasingly important. In this regard, it is of utmost importance to explain in detail the history of CDF and its key points An innovative team working method System engineering has features of both art and science since requires creativity and knowledge of systems engineers, but it also requires systems management and the application of a systematic disciplined approach. The traditional or the most classical design methodology is the sequential approach which means a sequence of specialists working in series. The overall design passes from a technical domain specialist, that works isolated from the other components of the design team, to another, during various design steps in successive time intervals. Lack of communication among the specialists can lead to wrong assumptions and obviously the main system parameters are not monitored in real-time. This method reduces the opportunity to find interdisciplinary solutions and to create system awareness in the specialists. An improved method is the centralized design, where the various technical domain specialists provide subsystem design information and data to a core team of one or more system engineers, but even this approach is not sequential. Concurrent Engineering is offered as an alternative to the classical approach and it provides better performance by taking full advantage of modern Information Technology (IT). Experts from various disciplines in the co-location could communicate in real-time and face to face. Since many disciplines are involved in the design process of complex systems, the concurrent approach has been proven particularly effective. Hence, the Concurrent Design Facility is a workspace and

46 information system allowing multidisciplinary experts working in a focused environment and conducting design collaboration History and status of CDF Some attempts on CE (Concurrent Engineering) began from 1980's in the field of aerospace and defense industry. A result of survey about CE was presented in 1993 by the Integrated Process Laboratory at the Concurrent Engineering Research Center (CERC), which was established at West Virginia University in 1988 by Defense Advanced Research Projects Agency (DARPA) to promote CE in United States industry. The results showed several advantages such as the possibility to reduce the design costs and to improve product quality at once. This survey clearly indicated that the most pressing need was to foster a teamwork environment, and the greater leverage exists in teamwork and process improvement. According to literature study, the first CDF with full features, which named with the Project Design Center (PDC) was opened by the Jet Propulsion Laboratory (JPL) in June of 1994[7]. The PDC provides a facility with multiple rooms for design teams to be used to conduct concurrent engineering sessions. Aerospace Corporation has developed the process and the tools for CE almost at the same time and they had been successfully applied to several programs. Based on the experience of the Aerospace Corporation, the JPL contracted the Aerospace Corporation to develop CEM (Concurrent Engineering Methodology) processes and tools for PDC. The Concept Design Center (CDC) was developed by the Aerospace Corporation in 1997, to enhance the support to its customers by providing a process for bringing together the conceptual design capabilities and experts. In the European space industry, concurrent engineering was also applied in the spacecraft design from the beginning of 1990'. The first example is provided by the Satellite Design Office (SDO) at DASA/Astrium, with the cooperation of the System Engineering (SE) group at the Technical University of Munich. An experimental design facility, Concurrent Design Facility (CDF), was created at the ESA Research and Technology Centre (ESTEC) at the end of 1998 and used to perform the assessment of several missions. The CDF is an Integrated Design Environment (IDE) based on the concurrent engineering methodology. Up to now,

47 more than 20 CDFs have been established around the world. These CDFs scatter in United States, Germany, France, Italy, Switzerland, United Kingdom and Japan, and are owned by governments, industries and universities Applications, Benefits and key elements of CDF Concurrent design is primarily used at ESA to assess technical, programmatic and financial feasibility of future space missions and new spacecraft concepts. Additionally, the ESA CDF (see Figure 10) is also used for many other multidisciplinary applications, such as payload instrument preliminary design, System of System (SoS) architectures and space exploration scenarios. Figure 10: ESA CDF in session Since 1994, two research teams, team-x and team-i, had conducted concurrent engineering design for space mission and space instrument in PDC of JPL. Applications of modern information systems enabled fundamental improvements to the system engineering process through the use of real time concurrent engineering. Many design teams have demonstrated dramatic savings in time and money compared with the traditional process for space systems conceptual design. As

48 reported in literature[7], improvements in efficiency obtained by team-x and PDC are significant and it should be noted that a dramatic reduction in average time to prepare proposals and very significant decrease in cost per proposal is achieved. The ESA/ESTEC summarizes the key elements on which the CDF implementation has been based: process, multidisciplinary team, integrated design model, facility, and infrastructure. These elements are detailed below: Process It is a fact that the space system has many interdependencies between components. This implied that the definition and evolution of each component has an impact on other components and that any change will propagate through the system. Early assessment of the impact of changes is essential to ensure that the design process converges on an optimized solution. The process starts with a preparation phase in which some representatives of the engineering team (team leader, system engineer, and selected specialists) and of the customer meet to refine and formalize the mission requirements, to define the constraints, to identify design drivers, and to estimate the resources needed to achieve the study objectives. Then the study kick-off takes place and the design process starts. It is conducted in a number of sessions in which all specialists must participate. This is an iterative process that addresses all aspects of the system design in a quick and complete fashion. One key factor is the ability to conduct a process that is not dependent on the path followed. At any stage it must be possible to take advantage of alternative paths or use professional estimates to ensure that the process is not blocked by lack of data or lack of decisions; A multi-disciplinary team Human resources are the most important and crucial element. A fundamental part of the CE approach is to create a highly motivated multi-disciplinary team that performs the design work in real-time. The challenge, the novelty of the method, the collective approach, the co-operative environment, the intense and focused effort and a clear and short term goal are all essential elements that contribute to personal motivation. To work effectively, the team members had to accept to use a new method of working, co-operate, perform

49 design work and give answers in real-time, and contribute to team spirit. For each discipline a position is created within the facility and assigned to an expert in that particular technical domain. Each position is equipped with the necessary tools for design modeling, calculations and data exchange. The choice of disciplines involved depends on the level of detail required and on the specialization of the available expertise. On the other hand, the number of disciplines has to be limited, especially in the first experimental study, to avoid extended debate and to allow a fast turn-around of design iterations; An Integrated Data Model The design process is model-driven using information derived from the collection and integration of the tools used by each specialist for his or her domain. A parametric model-based approach allows generic models of various mission/technological scenarios to be characterized for the study to perform. A parametric approach supports fast modification and analysis of new scenarios, which is essential for the real-time process. It acts as means to establish and fix the ground rules of the design and to formalize the responsibility boundaries of each domain. Once a specific model is established it is used to refine the design and to introduce further levels of detail. Each model consists of an input, output, calculation and results area. The input and output areas are used to exchange parameters with the rest of the system (i.e. other internal and external tools and models). The calculation area contains equations and specification data for different technologies in order to perform the actual modeling process. The results area contains a summary of the numeric results of the specific design to be used for presentation during the design process and as part of the report at the end of the study; An Appropriate Facility The team of specialists meets in the Concurrent Design Facility to conduct design sessions. The accommodation generally comprises a design room, a meeting room and project-support office space. The equipment location and the layout of the CDF are design to facilitate the design process, the interaction, the co-operation and the involvement of the specialists. The

50 facility is equipped with computer workstations dedicated to each technical discipline. A multimedia wall, supporting two or three large projector screens, is located in order to be visible from each working station. Each screen can show the display of each workstation, so that the specialists can present and compare design options or proposals and highlight any implications imposed on, or by, other domains; A Software Infrastructure An infrastructure to implement the Concurrent Design Facility outlined above requires tools for the generation of the model, integration of the domain models with a means to propagate data between models in real time, a means to incorporate domain-specific tools for modeling and/or complex calculations, a documentation-support system, and storage capability. The infrastructure must allow its users to work remotely from other facilities, and exchange information easily between the normal office working environment and the facility environment. Regarding the system model, Microsoft Excel spreadsheets are usually chosen for their availability and flexibility. The distribution of the model requires a mechanism to exchange relevant data between domains. This can be solved preparing a shared workbook to integrate the data to be exchanged, with macros to handle the propagation of new data in a controlled way. In some specific cases it can be more convenient not to use centralized data exchange, but rather to create a direct interface between those applications, such as the transfer of geometrical 3-D data of spacecraft-configuration to the simulation system[8]

51 Chapter 3 CIRA CONCURRENT DESIGN FACILITY 3.1 CIRA CDF FOR SPACE PROPULSION A Concurrent Design Facility for Space Propulsion is under development at CIRA in the frame of HYPROB program. The CDF will exploits concurrent engineering methodology to perform effective, fast and cheap liquid rocket engine (LRE) design. This discussion aims to describes the phases of design, the modules and their interaction. A liquid space propulsion system can be divided in the following main subsystems: Feeding (Tanks & Turbo-pumps), Cooling System, Injection System and Thrust Chamber (Combustion Chamber & Nozzle). Engineering Software shall be therefore composed by different modules dedicated to each subsystem, and by an upper level architecture module, that will compute the preliminary configuration of the LRE, based on main requirements, and shall allow the correct data exchange between the subsystem modules. Once a preliminary configuration is defined, more detailed verification programs will be used to verify the fulfilment of the requirements, including numerical tools for CFD computation and thermo-structural analysis. In Figure 11 a schematic view of a liquid space propulsion engine is shown

52 Figure 11: Liquid space propulsion engine - Schematic view In order to apply the logic discussed above, the following specialists are foreseen: Figure 12: CIRA Concurrent Design Facility for space propulsion - Specialist For each specialist a proper domain is foreseen. In addition, the following domains should be added: customer, team leader, CAD and schedule and planning. Figure 13 shows the total set of domains. Space propulsion architectural designer Thrust Chamber analyst Feeding/Turbopumps designer/analyst Cooling system designer/analyst Thermostructural analyst Computational Fluid Dynamics (CFD) Customer Team Leader Architecture (ARCH+ECOSIMPRO) Thrust Chamber (TCHA+INJE+ROCCID) Feeding System (FEED) Cooling System (COOL) Thermostructures (FEM) CFD Schedule and planning CAD Figure 13: CIRA Concurrent Design Facility for space propulsion - Domains

53 The process definition is one of the key points in a concurrent engineering approach. It is identified by three main phases: Requirements definitions; Subsystem preliminary design; Verification of requirements. For each module, the correct input/output parameters must be defined, along with the local parameters that are not exchanged with the other modules but are used within a single module for the subsystem sizing. Hereinafter the main steps of the overall process are briefly described. It must be underlined that, in the first version of the developed code, only pressure fed systems and expander cycle will be taken into account. Future developments will take into account other kind of cycles such as gas generator and staged combustion. Step 1 Requirements analysis First of all the team leader will analyze the requirements, typically thrust T (in the first version of the code) or I sp or ΔV and external pressure/exit conditions p e. Additional requirements, like propellants, cycle, total mass, can also be provided by the client; alternatively they are preliminarily defined within this step. Step 2 Preliminary architecture Architecture module (ARCH) receives the inputs from step 1 and calculates the O/F that maximizes the I sp, mass flow rate of oxidizer and fuel, preliminary geometry and the chamber pressure p c by using the engineering methods described in next sections. Architecture module activates then the subsystem modules. Step 3 Thrust chamber parameters TCHA receives the preliminary geometry, mass flow rates and chamber pressure as input and evaluates the wall heat flux q along the combustion chamber and nozzle, that is provided to the COOL module. Moreover, this module calculates the chemical 1-D composition and thermo-fluidynamic properties along the thrust chamber axis. Step 4 Injector and spray parameters

54 INJE receives in input the preliminary geometry, mass flow rates and chamber pressure and evaluates the pressure drop in the injection system, along with some typical parameters of the spray, like the breaking length and spray angle. Step 5 Cooling system Based on the inputs determined in the previous steps, COOL module calculates the number and geometries of the cooling channels, wall chamber thickness and the thermofluidynamic behavior of the coolant inside the channels. The necessary inputs depend on the chosen cycle. For pressure fed, regenerative cycle, case the following inputs are necessary: ARCH Preliminary geometry, mass flow rates and chamber pressure; TCHA Wall heat flux along the thrust chamber; INJE Pressure drop at the injection system (pressure at the injector inlet); For an expander cycle, the following inputs are required ARCH Preliminary geometry, mass flow rates and chamber pressure; TCHA Wall heat flux along the thrust chamber; FEED Pressure at the turbine inlet. FEM module will receive the same inputs and proceeds with the choice of the thrust chamber materials. Step 6 Feeding system FEED will receive from ARCH the general engine architecture; as previously clarified, in the first version of the code only pressure fed systems and expander cycle will be taken into account. If the architecture is a pressure fed cycle, FEED module will calculate the tank size and the feeding system layout based on the fuel conditions that must be realized at the inlet of the cooling system (from COOL module) and on the oxidizer conditions

55 at the injectors inlet (from INJE module). For an expander cycle the module will deal with the following subsystems: Fuel Pump: FEED will calculate the main pump parameters with the goal to obtain the required fuel conditions at the inlet of the cooling system, according to the COOL module; Oxidizer pump: FEED will calculate the main pump parameters with the goal to obtain the required oxidizer conditions at the injectors inlet, according to the INJE module; Turbine: FEED will receive the conditions from COOL module and calculate the turbine parameters. Once the general architecture and the main parameters of the subsystems have been defined, the detailed verification of the requirements can begin: EcosimPro SW will be used to simulate simultaneously all subsystems considering the transient phases; CFD software will be used to verify the data calculated by TCHA module; FEM software will perform detailed thermostructural calculations; ROCCID will verify that no combustion instabilities occur, along with the margins. If some requirements are not verified, or an optimization is needed, a further design iteration will be performed by ARCH module restarting from step 1, according to the modifications indicated by the Verification phase. Figure 14 shows the process implemented by CIRA Concurrent Design Facility for space propulsion[9]

56 Figure 14: CIRA Concurrent Design Facility for space propulsion - The process As already reported, CIRA CDF for space propulsion has been developed for a preliminary design of a liquid rocket engine. Therefore, in order to explain in detail the Cooling system module, on which the present work of thesis has been focused, it is of utmost importance to discuss the ARCH module and TCHA module. These modules provide the inputs for COOL module which it is strongly conditioned by ARCH and TCHA parameters. Thus, next sections are focused on the modules that are logically before the cooling module. 3.2 ARCHITECTURE MODULE The aim of the architecture module is to define a baseline for the sizing of the engine. As stated in the previous section, the architecture module is an upper level module that will use the input reported in step 1 to compute the first scheme of the propulsion engine. Considering Figure 15, it can be noted that a thrust chamber is constituted by three major elements: the combustion chamber, the exhaust nozzle and the injector

57 Figure 15:Typical basic configuration of a thrust chamber The combustion chamber and the exhaust nozzle can be preliminary sized using analytical and semi-empirical formulas. The first step is fixing the design point starting from the required thrust, nozzle conditions and selected propellants. Thus, the user has to decide which nozzle condition has to be adopted. In particular, three cases have been implemented. The first is to consider the external pressure pe equal to 1 atm and thus to consider the engine as it would be designed to ground conditions; the second is to impose pe equal to 0.01 atm and so to consider the engine as it would be designed for extra-atmospheric conditions; the last is to fix two geometrical area ratios: the combustion chamber area over the throat area and the exit nozzle area over the throat area (Ac/At and Ae/At). In particular, for the last case, values of Ac/At=3 and Ae/At=70 have been firstly considered. It is important to note that the software has been developed in order to cope with possible changes of those two area ratios. Considering the nozzle condition and the mixture chosen a large set of specific impulse values (I sp ) is imported through a precompiled database. Those databases are parameterized considering a wide range of operating conditions in terms of p c and O. In particular, those databases are the outcomes of a nested analyses performed F using the RPA (Rocket Propulsion Analysis) software, which assume the chemical equilibrium composition of the mixture. In general, this software utilizes a set of input parameters: combustion chamber pressure, propellant combination, mixture ratio or oxidizer excess coefficient or mass fractions of each component, list of

58 components at standard conditions or at assigned temperature, assigned enthalpy. RPA calculates combustion equilibrium and the properties of the reaction products. Additionally if the nozzle exit pressure, or alternatively the nozzle area ratio, is defined together with a chamber contraction area or mass flux, then the conditions at the nozzle throat, nozzle exit and the theoretical rocket engine performance are determined as well. This tool hence only performs combustion calculations and estimates the thruster performance. The process continues with the user selection of the design point. This can be the trivial maximization of the I sp or the result of a different strategy. For instance if the user has to take into account possible constraints like unfeasible values of p c and O F ratio. Once the design point has been chosen, and the I sp is evaluated, the size of the thrust chamber can be defined following the process reported hereinafter. From the imposed thrust and the resulting specific impulse, it s trivial to obtain the mass flow rate of propellant, and hence the fuel and oxidizer mass flows are provide by Eq. (3.2-1), (3.2-2) and (3.2-3): m = F I sp g 0 (3.2-1) ; m fu = m O F +1 (3.2-2) ; m ox = m O F O (3.2-3) F +1 Knowing the values of chamber pressure, mass flow rate and characteristic velocity, the throat area can be evaluated with the formula reported below (3.2-4). The characteristic velocity indeed is a result of the nested analysis directly linked to the chosen I sp value. A t = c m p c (3.2-4) The nozzle definition is completed by choosing the convergent-divergent geometry that can be bell or conic shaped. From this choice it is possible to size the combustion chamber. In an early design procedure phase, the chamber diameter has been evaluated considering the ratio between the chamber area and the throat area (Ac/At) equal to 3. However, if the user has decided to fix the pe as nozzle condition,

59 this choice, widely reported in literature and used to minimize chamber pressure losses, has been upgraded by the following semi-empirical correlation (3.2-5). A Ch A t = 8D t (3.2-5) The volume V CC, defined as the sum between the combustion chamber volume and the nozzle convergent part, is evaluated through the simple relation (3.2-6). V CC = L A t (3.2-6) Where the default value of the characteristic length is set equal to 1m, but can be easily changed accordingly to experience and literature, see Table 2. Table 2: Number of Characteristic Lengths of typical propellant combinations The convergent nozzle length is then evaluated by Eq. (3.2-7): L conv = D C D t 2 tan θ conv (3.2-7) Where θ conv is an angle closely linked to the maximum slope of the nozzle. This value is set in order to control the maximum slope of the converging part. Indeed, the maximum slope is finally evaluated in order to be sure that is included between 20 and 45. The volume of the combustion chamber V C is then evaluated subtracting from V CC the volume of the converging nozzle. This quantity is estimated with the formula (3.2-8) reported below: V conv = ( π 3 )L conv (R 2 C + R 2 t + R C R t ) (3.2-8)

60 If we suppose the combustion chamber to be cylindrical, its length can be easily obtained. Indeed, the chamber contraction ratio, defined as the ratio between the chamber cross-sectional area and the throat area (Ac/At), gives us the chamber diameter and, subsequently, its length. The last step is the sizing of the divergent part of the nozzle. If the user has fixed the area ratios as nozzle condition, then the evaluation of the nozzle exit area is trivial, on the contrary, if the user has imposed the value of the exit pressure, then, the theoretical nozzle expansion ratio can be obtained from the relation (3.2-9) of an ideal gas flow through a rocket nozzle. ε = A e A t = ( γ+1 ) γ+1 ( pc pe ) γ γ+1 γ 1 (3.2-9) γ 1 [1 (p e pc ) γ ] If the divergent part of the nozzle is supposed to be cone-shaped, the design will follow the configuration sketched in Figure 16: Figure 16: Conical nozzle contour The nozzle throat section has the contour of a circular arc with a radius that can be proportionally expressed in terms of throat radius R t. Its default value is 0.5 Rt, but it can range between 0.5 Rt and 1.5 Rt depending on engine size and experience considerations. It s worth to note that lower curvature values imply smaller dimensions and can produce lower thermal loads but higher heat peaks. The divergent half-cone angle θ div, instead, varies between 12 and 18 and has to be set by the user. Thus, the length of the divergent part of the nozzle can be evaluated by Eq. (3.2-10). L div = R t( ε 1)+R(sin θ div 1) tan θ div (3.2-10)

61 The subsequent evaluation of the exit radius is trivial: R e = εr t. If the parabolic design approximation for the bell nozzle is chosen, the design process follows the guidelines summarized in Figure 17. Figure 17: Bell nozzle contour The upstream throat contour is circular with a default radius 0.5 times the throat radius, terminating at the geometric throat. Even in this case its value can be changed according to experience. The downstream throat radius is also circular with radius Rt. It joins smoothly at the geometric throat with the upstream radius and continues till the angle θ n is reached. This procedure locates the final coordinates of the bell nozzle and let the user complete the design by drawing a smooth, parabolic curve using the parabola equation (3.2-11): y = ax 2 + bx + c (3.2-11) This equation can be solved imposing three known conditions, that are: the starting point; the value of the derivative in that point, that is equal to tan θ n and the value of the derivative in the ending point of the parabola, that is equal to tan θ e

62 Figure 18: Initial and final parabolic angles versus desired nozzle expansion ratio for different percent bell lengths of an equivalent 15 conical nozzle The last remaining decision is the shape of the converging part of the nozzle. Two possibilities have been considered and implemented. The first one, called straight consists of two rounded joints and a straight segment connecting them. Obviously the straight segment is tangent to the arcs. The second exploits a cubic function (Eq ) to link the end of the CC to the throat. y = ax 3 + bx 2 + cx + d (3.2-12) The second case ensures continuous second derivative, this is of utmost importance for the heat flux analysis and protection. In particular the solution can be achieved imposing four known conditions, that are: the starting and ending points positions, null first derivative at the throat section and the null second derivative in the inflection point, located halfway between the starting and ending points[9]. The Figure 19 shows the convergent nozzle contours for straight and cubic solutions Figure 19: Convergent nozzle contours for straight and cubic solutions

63 A preliminary geometry of the thrust chamber is finally obtained and plotted. Geometrical and fluid-dynamic parameters, such as chamber pressure, nozzle expansion ratio, O/F ratio and the chosen propellants, are then considered as inputs for lower level modules. To better understand the result obtained by this routine, Figure 20 shows an example of the geometry obtained considering a conic shaped nozzle with a cubic convergent part. Figure 20: An example of the Thrust Chamber geometry evaluated by the ARCH module 3.3 THRUST CHAMBER MODULE Thrust chamber module is a point of connection among ARCH module and COOL module. Starting from the architecture module s output, such as the chamber pressure, the nozzle expansion ratio, the O/F ratio and the chosen propellants, the aim of the thrust chamber module is to perform an analysis of the combustion process, providing temperature and chemical composition along the axis of the thrust chamber. In order to accomplish this target, the simplifying assumption of chemical

64 equilibrium is considered valid. In particular the module uses the software named CEA (Chemical Equilibrium Analysis). CEA is a tool developed by Gordon and McBride at NASA Glenn/Lewis Research Center[4]. The name of this software suggests that the chemical equilibrium has been assumed in the combustion chamber. The chemical equilibrium of a reacting system permits to evaluate, in a simplified way, the theoretical thermodynamic properties that are useful for the design of several complex systems such as compressors, turbines, nozzles, engines. The equilibrium is usually described by either of two equivalent formulations, equilibrium or minimization of free energy. In particular, the minimization of free energy formulation is used in the CEA program. In the thrust chamber the propellants react to form hot gases. Hence, the thrust chamber module is focused on the hot gas properties. The hot gas characteristics are obtained in three locations, that are: combustion chamber, throat region and the nozzle exit. To obtain this result, a MATLAB routine writes the input file for the software CEA, then it externally runs the program and finally loads and saves the resulting data. The physical quantities that are particularly interesting in this phases are: Mach number, characteristic velocity, thrust coefficient, specific impulse, viscosity, thermal conductivity, Prandtl number, velocity, sound velocity, specific heat ratio, heat specific, temperature, pressure and Reynolds number. Most of those parameters have been used in the first step of the cooling module

65 Chapter 4 COOLING SYSTEM MODULE OF CIRA CDF 4.1 OVERVIEW The present chapter describes the main objective of the present work i.e., the cooling system module for CIRA CDF. In particular, this chapters describes in detail all key passages of numerical investigation. Obviously, this approach 1-D involves an approximate description of the phenomena, but the obtained results (see Chapter 5), are completely satisfactory for a phase 0/A of a space project. The procedure design follows step by step the engineering formulas reported in the literature. After the choice of the geometry of the thrust chamber and the operating condition performed by architecture module, several models for the cooling system design have been implemented. In particular, the coolant behavior changes with different sizes and number of the cooling channels that surround the thrust chamber. For the first development of the Cooling system module, the cross-sectional of the channels has been assumed circular. In addition to the pressure drop, temperature and other thermo-fluidynamics properties of the coolant may vary with a different formulations of a friction factor. Several thermo-fluidynamics characteristics of the coolant will be provided by a software tool, called CoolProp integrated with CDF Software. The primary objective of cooling is to preserve the chamber and nozzle walls from the huge heat flux coming from hot gasses passage in the inner part of the thrust chamber. The high temperatures may exceed 3600 K, the pressures may exceed 30 MPa and the heat fluxes can reach 100 MW/m 2. This represents a very challenging problem. Many cooling techniques have been developed in liquid rocket engine

66 manufacturing: regenerative cooling, radiation cooling, dump cooling, film cooling, transpiration cooling and ablative cooling. The first version of the engineering software will refer to a regenerative cooling architecture, therefore regenerative cooling technique will be hereinafter discussed. The basic concept (see Figure 21) is to use liquid propellant to cool the thrust chamber. The propellant flows inside the cooling channels, increasing its energy and changing phase from liquid to gas. Finally it is injected in the combustion chamber. Figure 21: Regenerative cooling architecture The cooling system module receives nozzle geometry by Architecure Module (ARCH) and several thermo-fluidynamics properties of hot gases by Thrust Chamber Module (TCHA). 4.2 COOLING SYSTEM MODULE Heat flux analyses The design of thrust-chamber cooling channels will start with the calculation of the heat transfer from combustion gasses, through the solid walls, to cooling channels. As a first step, a 1-D steady state condition is considered. The heat transfer from combustion gases through the wall to the coolant region (see Figure 22) can be expressed by the equation (4.2-1):

67 q = h g (T aw T wg ) = ( K t ) (T wg T wc ) = h c (T wc T co ) (4.2-1) Each part of this equation has to be modeled. Figure 22: Heat transfer for schematic regenerative cooling Let s start with the gas side convective heat flux reported in eq. (4.2-2) q = h g (T aw T wg ) (4.2-2) Obviously, the estimate of heat flux is preceded by some steps which provide some parameters. First of all, the designer must choose the value of T wg at throat. Indeed, the user has the possibility to set up the gas side wall temperature through a command menu. According to literature, the default value is imposed equal to 700 K, but this value can be easily changed to consider different throat conditions. Since the throat is the critical point, the wall temperature is greater in that point than in other region of the thrust chamber. Moreover, in along the nozzle, the hot gases temperature decreases by isentropic expansion law, hence the gas side wall temperature decreases. In preliminary design, the value of gas side wall temperature at the beginning of the chamber has been set equal to 2 T 3 wg while at exit of the nozzle this parameters has

68 been fixed at 1 3 T wg. In this way, the other parameters will be calculated in the same points. The T aw of the combustion gas may is obtained using eq. (4.2-3) r (γ T aw = (T c ) ns [ 2 ) M x 2 ] = (T γ 1 c ) nsr 1 + ( 2 ) M x 2 (4.2-3) Where r is the local recovery factor and represents the ratio of the frictional temperature increase to the increase caused by a adiabatic compression. Below, two simplified correlations based on Prandtl number, are reported. r = (Pr) 0.5 for laminar flow (4.2-4) r = (Pr) 0.33 for turbulent flow (4.2-5) Bartz proposed a semiempirical evaluation of the Nusselt numbers, and thus of the gas-side heat transfer coefficient h g [51] [5]. Nu = Re 0.8 Pr 0.3 (4.2-6) where Nu= h g d k Hereinafter, a modified Bartz equation is reported: (μ0.2 C p h g = [ 0.2 D t Pr 0.6 ) ( (p 0.8 c) ns g c ) ( Dt 0.1 ) ] ( A 0.9 t r ns t A ) σ (4.2-7) (4.2-8) The correction factor for property variation across the boundary layer is evaluated as specified in Eq. (4.2-9) σ = T wg [ 1 2 (1 + (T c ) ns 1 γ 1 2 M2 ) ] γ 1 (4.2-9) [1 + 2 M2 ] In this way, each of these magnitudes is known in three points, that are: in the combustion chamber, at throat and at the exit of the nozzle. But, for some of these

69 magnitudes, the total distribution along the thrust chamber must be known. Therefore, the variation of T wg along the axis of thrust chamber has been approximated by a logarithmic trend. The distribution of T aw and h g has been similarly constructed. Successively, the equation (4.2-2) has been used for calculate the heat flux along the thrust chamber (see Chapter 5 for results). This distribution has been utilized for a first evaluation of the phenomena, but it has been corrected after some considerations. Because in combustion chamber the chemical equilibrium has been assumed and the wetted area remains constant, the heat flux distribution along the chamber can be reasonably assumed constant. In the Typical axial heat transfer rate distribution for liquid propellant thrust chambers, the peak is always at the nozzle throat and the lowest value is usually near the nozzle exit. The procedure design continues with the choice of the wall thickness. The default value is fixed at m, but the user, according its experience can choose a more appropriate value. Of course, this choice will be the result of a compromise among rocket weight, performance and manufacturing constrains. However, wall thickness can be also evaluated by the relation (4.2-10)[10]: t = f s p c D c 2(σ y ξ 0.6p c ) (4.2-10) Where p c is the chamber pressure, D t is the throat diameter, σ y is the yield stress that depends by material and temperature, ξ is the joint coefficient and f s is the safety factor. Similarly, the user will impose the distance between channels at throat section. The default value, according to literature studies, is fixed at m. Next step is dedicated to the evaluation of the thermal conductivity of the thrust chamber walls. It depends on the selected material and it changes with temperature. Hence, the designer must choose the walls material. CIRA CDF SW is already provided by the following material database: CuCrZr alloy, copper, gold, aluminium, iron, niobium. If the material desired it not yet schematized, the user can

70 set up a linear or constant law of thermal conductivity. The Figure 23 shows the thermal conductivity distribution of these materials[11]. Figure 23: Variation of thermal conductivity with temperature for typical metallic elements and alloy At this point, exploiting the second part of the semplified Fourier equation (4.2-11) for 1D - linear assumption, T wc can be evaluated : T wc = T wg t q k (4.2-11) Cooling Channels geometry The process design continues with the evaluation of the cooling channels geometry. As already told, considering an early simplified approach, the cross-section area of

71 the cooling channels has been assumed circular. Figure 24 shows a sketch of the channel. Figure 24: Detail view cooling channel geometry In regenerative cooling process, the coolant, generally the fuel enters passages at the nozzle exit of the thrust chamber nozzle. Thus, the coolant passes through the throat region and reaches the exit near the injection plane. This path is represented in the Figure 25. Figure 25: Cross-sectional view of a regenerative cooling thrust chamber showing the flows directions The nozzle throat region usually experiences the highest heat flux and therefore is the most difficult to cool down. For this reason the first cooling passages section are designed in such a way that the coolant velocity is highest at the critical regions. This is achieved considering the minimum cross-section area of the coolant passage at nozzle throat. As show in Figure 26, the cross-section area of cooling passages scales according to the region of thrust chamber to uniformly cool the entire wetted area

72 Figure 26: Typical cross-sectional scaling of a cooling channels along axial direction The geometrical sizing of channels starts with the imposition of the diameter at throat and proceeds through the entire thrust chamber. The user can consider several different values at throat, in order to understand the influence of the diameter and finally choose the best solution. As already told, the diameter changes along the axial direction according to the region of thrust chamber. Hence, in the throat region the diameter will be the tiniest. For manufacturing reason, the diameter of channels can t be less than 0.8 mm. The user can therefore analyze several solution obtained considering different values of the cooling channels diameter. The number of cooling channels that surround the thrust chamber can be easily calculated when some parameters have been considered. From Eq. (4.2-12) it is possible to denote that the number of channels depends by some factors such as: the geometry of the nozzle, wall thickness t and distance between coolant passages s

73 n = 2π[(t+r t )+(d ch ( r e rt ))] [(d ch +s)( r e rt )] (4.2-12) For every selected size of the channels diameter, a different number of cooling channel is obtained. In this regard, a study has been performed to evaluate the relationship between the size of the cooling channel and the variations in thermofluidynamic properties of the coolant. Those effects will be widely reported in the Chapter 5. From n, using Eq. (4.2-13) it is trivial to obtain the fuel flow rate of the single channel: m ch fu = m fu n (4.2-13) Before assessing the thermo-fluydinamic characteristics of the coolant, the wetted area can be calculated from simple geometric considerations. Considering the thrust chamber contour, the wetted area, useful for further calculations, can be easily evaluated. This is the surface surrounding the hot gas flow and will be necessary to obtain thermo-fluydinamic parameters of the coolant flow through the channels Coolant flow analysis This paragraph describes the coolant thermo-fluidynamic properties evaluation of the coolant, once that the geometry of the channels has been designed as described in the previous paragraph. The developed routine is based on a software tool called CoolProp[12] hereinafter described. For what concern pressure losses, different models of friction factor have been implemented and will be described in this chapter. Final results will be presented in Chapter

74 CoolProp The numerical procedure developed in the present thesis, is supported by CoolProp[12] that allows the evaluation of the thermo-fluydinamics behavior of the coolant. CoolProp is a C++ library that implements: Pure and pseudo-pure fluid equations of state and transport properties for 114 components; Mixture properties using high-accuracy Helmholtz energy formulations (or cubic EOS); Correlations of properties of incompressible fluids and brines; Highest accuracy psychrometric routines CoolProp is based on Helmholtz energy formulations and all thermodynamic properties of interest can be obtained directly from partial derivatives of the Helmholtz energy. It should be noted that the EOS are typically valid over the entire range of the fluid, from subcooled liquid to superheated vapor, to supercritical fluid. Cooling System tool numerical description The first step of the developed procedure relies on the reading of input parameters from the feeding module, therefore the thermofluidynamic state of the fluid at the beginning of the channel is known. Coolprop uses initial temperature and pressure. A series of magnitudes, reported hereinafter, are calculated by CoolProp: thermal conductivity, density, specific heat at constant pressure, sound velocity, viscosity and the phase. During the numerical implementation, an issue occurs when the coolant temperature approaches at one of the transition phase (as example form K to K for methane). In particular, the temperature value has been set equal to a value immediately greater than supercritical value in order to avoid the well-known strong variation of thermodynamic coefficients. In the common range of pressure occurring in an LRE liquid Hydrogen is in supercritical conditions, while Methane usually works in transcritical conditions (see Appendix A for details)

75 The design procedure continues with the evaluation of other parameters such as Reynolds number and Prandtl number as shown in the Eq. (4.2-14) and Eq. (4.2-15): Re = ρvd μ Pr = C pμ k (4.2-14) (4.2-15) Moreover, by Eq. (4.2-16) the coolant velocity in the channels can be calculated: v = m fu ch A ch ρ (4.2-16) In supercritical/transcritical conditions the heat is transferred through a vapor-film boundary layer and the coolant-side heat-transfer coefficient can be estimated from one of the following equation (4.2-17) [13] and (4.2-18)[4]: h c = ( k d ) (Re0.8 Pr 0.4 ) ( T 0.1 co ) T wc (4.2-17) h c = ( k d ) (Re0.8 Pr 0.4 ) ( T co T wc ) (4.2-18) Both the equation will be utilized in first evaluation, but the (4.2-17) will be considered in further verifications. Of course, each part (i.e., gas side fluidynamic, wall thermal and coolant side fluidynamic analyses) shall be approached by means of deeper methods, but the scope of the present paragraph is to underline the basic approach in a phase A of design procedure. The next step focuses on the evaluation of the global coefficient of heat transfer by Eq. (4.2-19) [12]: H = ( 1 h g + t k + 1 h c ) (4.2-19) This parameter is crucial in the evaluation of coolant temperature, it depends, substantially, by hot gases properties, thrust chamber geometry, chamber wall thickness, wall material and coolant properties

76 In the first implementation of the SW, the global coefficient hasn t been used for the evaluation of the heat flux and thus of the coolant temperature. The heat flux has been simply obtained using the hot gas side heat flux coefficient, resulting in a less accurate estimation of temperature increase of the coolant Indeed, the variation of the coolant temperature (T 2co T 1co ) along the channels can be calculated by means following correlation: T 2co = T 1co + A w q m fu C p (4.2-20) It depends, substantially, by heat flux, wetted area, coolant mass flow rate and coolant specific heat. Successively, a new evolution of the coolant temperature, has been evaluated by means of a deeper evaluation of the heat flux, using the global coefficient approach (see equation ): q = H(T aw T co ) (4.2-21) Then, the coolant temperature is calculated by means of eq. (4.2-20) This step is one of most important because the coolant temperature is an input for CoolProp software and the other thermos-fluidynamic properties will be strongly affected by it. The coolant absorbs heat flux along the channels and its temperature increases. In this way, the coolant will be injected in combustion chamber with a greater level of enthalpy. However, this phenomena doesn t improve significantly the rocket engine performance in terms of efficiency. The last step is focused on the coolant pressure along the channels. As already told, the pressure is the other input for CoolProp. From the literature [1] [5] the pressure drop can be calculated by the following equation: Δp = f L D H 1 2 ρv2 (4.2-22)

77 Pressure drop depends by density, velocity and geometric considerations. In particular, the coolant behavior in terms of pressure drop is strongly conditioned by the friction factor. In this work, several formulations of friction factor have been considered and implemented: For smooth pipes with Re < 10 6 Zandbergen proposes to use the following relationships by Poisseuille and Blausius[4]: f = 64 Re ( 1 Re )0.25 Re < < Re < (4.2-23) { ( 1 Re ) < Re < For non-smooth pipes with Re > 10 6, Zandbergen proposes using the following relation by Nikuradse[4]: f = 8 ( log ( e d )) 2 (4.2-24) The Moody diagram. To each diameter channel length, a given value of e d is provided by a particular curve on the diagram that is function of Reynolds number (see Figure 27 )

78 Figure 27: Moody Diagram Colebrook relation[14]: f = 0.25 [log( e 3.7D Re 0.9 )]2 (4.2-25) These models of a friction factor have been implemented, however, another relationship has been considered for the pressure drop estimation[15]: Δp = f L D H 1 2 ρv2 + ρ 1 v 1 (v 2 v 1 ) (4.2-26) The only differences between Eq. (4.2-22) and Eq. (4.2-26) consists in the term ρ 1 v 1 (v 2 v 1 ). In the next chapter its effect will be highlighted and plotted along the channel direction. The presented mathematical formulations has been used and compared to describe the coolant behavior along the axis of the channels

79 Chapter 5 RESULTS 5.1 Overview The present chapter provides the description of the validation of the design cycle focusing on the developed tool described in chapter 4. The chosen test case is the CIRA demo LOX/CH4 described in [16] and [17]. Moreover a parametric analysis varying friction factor, cooling channel diameter size and heat flux evaluation will be presented. The first step is to run the architecture module and fix the design point starting from the required thrust, nozzle conditions and selected propellants. Then, the user has to decide which nozzle (bell conic, etc.) has to be adopted. In the present work a conic-cubic nozzle has been chosen, the value of thrust has been fixed equal to N, the chamber pressure has been set equal to 55 bar (common value for this class of Rocket Engine), the convergence angle and divergence angle have been imposed equal to 22.5 and 20 respectively. For the nozzle conditions, the first strategy has been adopted, hence p e = 1 atm. Liquid oxygen and liquid methane have been chosen as propellants and their mixture ratio has been fixed at 3.4 maximizing specific Impulse (see [16] and [17]). Finally the characteristics length has been fixed equal to 0.94 m, while the throat radius ratio will be equal to 1.5. This procedure design provided a geometrical profile of thrust chamber as shown in Figure 28:

80 y [m] x [m] Figure 28: Geometrical profile of thrust chamber The inlet conditions for the cooling channels of the methane have been fixed for temperature and pressure. In particular at inlet of channels T co has been imposed equal to 110 K, while the pressure equal to 160 bar. A first evaluation has been realized considering a only diameter of channels at throat. 5.2 Validation of the cooling system design cycle In this paragraph a validation of the performed work will be shown. This final result has been obtained comparing the simulation performed using the developed module with the results obtained by the 3 ton class LOX/CH4 LRE developed at CIRA in the framework of HYPROB Program. The HPRB-BREAD project has been defined in order to develop and test a LOX/LCH4 rocket engine regeneratively cooled ground demonstrator and related Breadboards for technology and design validation. The architecture considered for the demonstrator, in line with the project key level requirements, is a regenerative cooled thrust chamber for ground testing. In the HYPROB-demonstrator a counter-flow architecture will be considered for the

81 chamber cooling system (see Figure 29 and Figure 30). In this type of architecture, the coolant (LCH4) is injected liquid into the fuel manifold and enters the cooling jacket counter flow with respect to the combustion gases. After being heated it is injected directly in the fuel dome and then from the injector in the combustion chamber where mixes, atomizes and burns with liquid oxygen. Figure 29: Architecture concept Figure 30 Counter flow architecture for the cooling jacket The main CIRA demo parameters are shown in the following Table 3:

82 Parameter D t [m] Throat diameter D c [m] Combustion chamber diameter D e [m] Exit diameter V c [m 3 ] Combustion chamber volume 2.693e-3 L cham [m] Combustion chamber length L conv [m] Convergent nozzle length L div [m] Divergent nozzle length n Channels number 96 AR Aspect Ratio 1.4 Convergence angle 22.5 Divergence angle 20 Table 3: Main geometric parameters of HYPROB-demonstrator With those parameters, the geometric configuration of the thrust chamber is represented in Figure 31: Figure 31: Geometric profile of thrust chamber The geometric profile used for the performed work shows small differences with respect to the thrust chamber configuration shown in Figure 31: L div is slighty smaller than L div of the demonstrator (see Figure 28). Moreover, the channels of

83 HYPROB-demonstrator has been developed with a rectangular shape (see Figure 32), but in the present work the numerical simulations have been carried out considering circular cooling channels (see Figure 26). This difference involves a similar number of channels: 96 for the demonstrator and 94 for the numerically simulated module. Of course, in this way the comparison between two models is reasonable, because the differences are very small. For both models, CuCrZr alloy has been chosen for inner wall material. Table 4 shows performance parameters of HYPROB-demonstrator. Figure 32: Cooling system channel and brazing interface Performance parameters c [m/s] Characteristic velocity p c [Pa] Chamber pressure T c [K] Chamber temperature 3542 I sp [s] Specific Impulse C f Thrust coefficient F (sea level) [N] Thrust T CH4 [K] Inlet coolant temperature 110 m fu [kg/s] Mass flow rate (fuel) Mixture ratio 3.4 Table 4: Main performance parameters of HYPROB-demonstrator