This is predominantly a qualitative study on the manufacturing of micronozzles with an additive manufacturing (AM) technique, namely the laser-powered powder bed fusion (PBF-LB).  Manufacturing of micronozzles with standard microelectromechanical system technology often results in 2.5-D or close to 3-D structures and does not yield a fully rotationally symmetric nozzle. For this reason, AM can be a better solution. However, the structures obtained with PBF-LB exhibit very rough surfaces which will impair the performance of the micronozzle. To improve the surface finish electropolishing was performed on the interior walls.  Given the shape and the scale of the components, uniformity of the polishing is a challenge, calling for an inventive electrode configuration and electrolyte feed solution. The approach was to integrate an electrode on the inside of the converging part of the nozzle, to serve as a cathode for the electropolishing, already in the process, and to make the nozzle itself the vital part of the fluidic system.  With this, titanium micronozzles were manufactured with throat diameters varying between 300 and 800 μm. With the resolution of the used AM technique, it was possible to integrate the internal electrode in the micronozzles with a designed throat diameter down to 600 μm. Below this, the anode, and cathode, sometimes made contact short-circuiting the cell. Profilometry showed a decrease of the average surface roughness (𝑅𝑅𝑎𝑎) with 15-60 % for the electropolished micronozzles. The Schlieren imaging showed an exhaust that followed the throat’s axial direction and also demonstrated pressure disks and, hence, a supersonic jet exhaust. This study has shown that AM is a viable choice for manufacturing of rotationally symmetric micronozzles, and that electropolishing could be used to decrease the surface roughness on their inside uniformly with the integration of a cathode.

No description available.

The increasing demand for higher performance in rocketlaunchers promotes the development of nozzles with higherperformance, which basically is achieved by increasing theexpansion ratio. However, this may lead to flow separation andensuing instationary, asymmetric forces, so-called side-loads,which may present life-limiting constraints on both the nozzleitself and other engine components. Substantial gains can bemade in the engine performance if this problem can be overcome,and hence different methods of separation control have beensuggested. However, none has so far been implemented in fullscale, due to the uncertainties involved in modeling andpredicting the flow phenomena involved. In the present work the causes of unsteady and unsymmetricalflow separation and resulting side-loads in rocket enginenozzles are investigated. This involves the use of acombination of analytical, numerical and experimental methods,which all are presented in the thesis. A main part of the workis based on sub-scale testing of model nozzles operated withair. Hence, aspects on how to design sub-scale models that areable to capture the relevant physics of full-scale rocketengine nozzles are highlighted. Scaling laws like thosepresented in here are indispensable for extracting side-loadcorrelations from sub-scale tests and applying them tofull-scale nozzles. Three main types of side-load mechanisms have been observedin the test campaigns, due to: (i) intermittent and randompressure fluctuations, (ii) transition in separation patternand (iii) aeroelastic coupling. All these three types aredescribed and exemplified by test results together withanalysis. A comprehensive, up-to-date review of supersonic flowseparation and side-loads in internal nozzle flows is givenwith an in-depth discussion of different approaches forpredicting the phenomena. This includes methods for predictingshock-induced separation, models for predicting side-loadlevels and aeroelastic coupling effects. Examples are presentedto illustrate the status of various methods, and theiradvantages and shortcomings are discussed. A major part of the thesis focus on the fundamentalshock-wave turbulent boundary layer interaction (SWTBLI) and aphysical description of the phenomenon is given. Thisdescription is based on theoretical concepts, computationalresults and experimental observation, where, however, emphasisis placed on the rocket-engineering perspective. This workconnects the industrial development of rocket engine nozzles tothe fundamental research of the SWTBLI phenomenon and shows howthese research results can be utilized in real applications.The thesis is concluded with remarks on active and passive flowcontrol in rocket nozzles and directions of futureresearch. The present work was performed at VAC's Space PropulsionDivision within the framework of European spacecooperation. Keywords:turbulent, boundary layer, shock wave,interaction, overexpanded,rocket nozzle, flow separation,control, side-load, experiments, models, review.

<p>The increasing demand for higher performance in rocketlaunchers promotes the development of nozzles with higherperformance, which basically is achieved by increasing theexpansion ratio. However, this may lead to flow separation andensuing instationary, asymmetric forces, so-called side-loads,which may present life-limiting constraints on both the nozzleitself and other engine components. Substantial gains can bemade in the engine performance if this problem can be overcome,and hence different methods of separation control have beensuggested. However, none has so far been implemented in fullscale, due to the uncertainties involved in modeling andpredicting the flow phenomena involved.</p><p>In the present work the causes of unsteady and unsymmetricalflow separation and resulting side-loads in rocket enginenozzles are investigated. This involves the use of acombination of analytical, numerical and experimental methods,which all are presented in the thesis. A main part of the workis based on sub-scale testing of model nozzles operated withair. Hence, aspects on how to design sub-scale models that areable to capture the relevant physics of full-scale rocketengine nozzles are highlighted. Scaling laws like thosepresented in here are indispensable for extracting side-loadcorrelations from sub-scale tests and applying them tofull-scale nozzles.</p><p>Three main types of side-load mechanisms have been observedin the test campaigns, due to: (i) intermittent and randompressure fluctuations, (ii) transition in separation patternand (iii) aeroelastic coupling. All these three types aredescribed and exemplified by test results together withanalysis. A comprehensive, up-to-date review of supersonic flowseparation and side-loads in internal nozzle flows is givenwith an in-depth discussion of different approaches forpredicting the phenomena. This includes methods for predictingshock-induced separation, models for predicting side-loadlevels and aeroelastic coupling effects. Examples are presentedto illustrate the status of various methods, and theiradvantages and shortcomings are discussed.</p><p>A major part of the thesis focus on the fundamentalshock-wave turbulent boundary layer interaction (SWTBLI) and aphysical description of the phenomenon is given. Thisdescription is based on theoretical concepts, computationalresults and experimental observation, where, however, emphasisis placed on the rocket-engineering perspective. This workconnects the industrial development of rocket engine nozzles tothe fundamental research of the SWTBLI phenomenon and shows howthese research results can be utilized in real applications.The thesis is concluded with remarks on active and passive flowcontrol in rocket nozzles and directions of futureresearch.</p><p>The present work was performed at VAC's Space PropulsionDivision within the framework of European spacecooperation.</p><p><b>Keywords:</b>turbulent, boundary layer, shock wave,interaction, overexpanded,rocket nozzle, flow separation,control, side-load, experiments, models, review.</p>

NR 20140805

There has recently been an increased interest in methane as a rocket propellant due to its physical properties as well as the possibility of in-situ resource utilization in places like Mars. As part of ESA’s Future Launcher Preparatory Program, KTH in cooperation with GKN Aerospace has started the MERiT program, which seeks to study the characteristics of methane under conditions found in rocket nozzle cooling channels. In particular, the current work examines the influence of different wall temperatures, fluid flow rates, and fluid residence times on methane pyrolysis due to the catalytic properties of nickel based metals. Pyrolysis is the thermo-catalytic decomposition of methane, which results in the creation of hydrogen and solid carbon in the cooling channels. This can affect the performance of the rocket engine, the cooling channels, as well as the lifespan of the engine, which makes the process important to quantify when designing highly reusable engines. A chemical kinetics computer model has been developed, which has been used to quantify the most important parameters for methane pyrolysis. Based on these results, a small-scale pyrolysis experimental setup has been developed and used to characterise methane pyrolysis for different material temperatures and gas flow rates. The experimental setup has been proven to work and consistently provide pyrolysis at temperatures between 600 ◦C to 700 ◦C, although more work on the data collection side, in particular with regards to a gas chromatograph and a scanning electron microscope, is required to quantify and compare different experiments.

Conjugate heat transfer analysis has been carried out on an 89kN thrust chamber in order to evaluate whether combined discrete film cooling and regenerative cooling in a rocket nozzle is feasible. Several cooling configurations were tested against a baseline design of regenerative cooling only. New designs include combined cooling channels with one row of discrete film cooling holes near the throat of the nozzle, and turbulated cooling channels combined with a row of discrete film cooling holes. Blowing ratio and channel mass flow rate were both varied for each design. The effectiveness of each configuration was measured via the maximum hot gas-side nozzle wall temperature, which can be correlated to number of cycles to failure. A target maximum temperature of 613K was chosen. Combined film and regenerative cooling, when compared to the baseline regenerative cooling, reduced the hot gas side wall temperature from 667K to 638K. After adding turbulators to the cooling channels, combined film and regenerative cooling reduced the temperature to 592K. Analysis shows that combined regenerative and film cooling is feasible with significant consequences, however further improvements are possible with the use of turbulators in the regenerative cooling channels.

La vectorisation de la poussée d'une tuyère propulsive supersonique axisymétrique est étudiée par le biais d'une injection fluidique transversale dans sa partie divergente. Cette étude menée dans le cadre du programme PERSEUS du CNES a été motivée par la recherche d'une solution alternative au pilotage conventionnel de la poussée par actionneurs mécaniques. Le travail de la thèse, tout en s'appuyant sur des approches expérimentale et numérique, comprend essentiellement une large étude paramétrique concernant principalement la position de l'injection, la forme de la tuyère, la nature et le débit du fluide injecté. L'analyse des résultats montre que pour certaines configurations optimales, des angles de déviation pertinents peuvent être obtenus pour des taux d'injections modérés. L'analyse numérique étendue aux écoulements chauds multi-espèces, plus proches des applications réelles, a montré que la vectorisation fluidique reste très performante lors de l'injection de produits de combustion dans le divergent. / Secondary injection into the divergent section of a supersonic rocket nozzle is investigated for the fluidic thrust vectoring effects. The study was conducted in the framework of CNES PERSEUS program and was motivated by the need for an alternative vectoring solution aimed for a small space launcher. The thesis work, based on the combined experimental and numerical approaches, essentially comprises of a wide parametric study mainly concerning the position of the injection, the shape of the primary and injection nozzles, flow regime, gas thermophysical properties and injected fluid mass-flow-rate. The analysis shows that for some optimal configurations, pertinent deflection angles can be obtained using the moderate injection rates. Furthermore, the study was extended to the hot flow multi-species investigation, simulating a case closer to the real applications. This numerical analysis indicated that the fluidic vectoring method remains effective with injection of combustion products into the divergent section of a propulsive rocket nozzle.

Convergent-Divergent (C-D) nozzles are used in rocket engines to produce thrust as a reaction to the acceleration of hot combustion chamber gases in the opposite direction. To maximize the engine performance at high altitudes, large area ratio, bell-shaped or contoured nozzles are used. At lower altitudes, the exit pressure of these nozzles is lower than the ambient pressure. During this over-expanded condition, the nozzle-internal flow adapts to the ambient pressure through an oblique shock. But the boundary layer inside the divergent portion of the nozzle is unable to withstand the pressure rise associated with the shock, and consequently flow separation is induced. Numerical simulation of separated flows in rocket nozzles is challenging because the existing turbulence models are unable to correctly predict shock-induced flow separation. The present thesis addresses this problem. Axisymmetric, steady-state, Reynolds-Averaged Navier-Stokes (RANS) simulations of a conical nozzle and three sub-scale contoured nozzles were carried out to numerically predict flow separation in over-expanded rocket nozzles at different nozzle pressure ratios (NPR). The conical nozzle is the JPL 45◦-15◦ and the contoured nozzles are the VAC-S1, the DLR-PAR and the VAC-S6-short. The commercial CFD code ANSYS FLUENT 13 was first validated for simulation of separated cold gas flows in the VAC-S1 nozzle. Some modeling issues in the numerical simulations of flow separation in rocket nozzles were determined. It is recognized that compressibility correction, nozzle-lip thickness and upstream-extension of the external domain are the sources of uncertainty, besides turbulence modeling. In high-speed turbulent flows, compressibility is known to affect dissipation rate of turbulence kinetic energy. As a consequence, a reduction in the spreading rate of supersonic mixing layers occurs. Whereas, the standard turbulence models are developed and calibrated for incompressible flows and hence, do not account for this effect. ANSYS FLUENT uses the compressibility correction proposed by Wilcox [1] which modifies the turbulence dissipation terms based on turbulent Mach number. This, as shown in this thesis, may not be appropriate to the prediction of flow separation in rocket nozzles. Simulation results of the standard SST model, with and without the compressibility correction, are compared with the experimental data at NPR=22 for the DLR-PAR nozzle. Compressibility correction is found to cause under-prediction of separation location and hence its use in the prediction of flow separation is not recommended. In the literature, computational domains for the simulation of DLR subscale nozzles have thick nozzle-lips whereas for the VAC subscale nozzles they have no nozzle-lip. Effect of nozzle-lip thickness on flow separation is studied in the DLR-PAR nozzle by varying its nozzle-lip thickness. It is found that nozzle-lip thickness significantly influences both separation location and post-separation pressure recovery by means of the recirculation bubbles formed at the nozzle-lip. Usually, experimental values of free stream turbulence are unknown. So conventionally, to minimize solution dependence on the boundary conditions specified for the ambient flow, the computational domain external to the nozzle is extended in the upstream direction. Its effect on flow separation is studied in the DLR-PAR nozzle through simulations conducted with and without this domain extension. No considerable effect on separation location and pressure recovery is found. The two eddy-viscosity based turbulence models, Spalart-Allmaras (SA) model and Shear Stress Transport (SST) model, are well known to predict separation location better than other eddy-viscosity models, but with moderate success. Their performances, in terms of predicting separation location and post-separation wall pressure distribution, were compared with each other and evaluated against experimental data for the conical and two contoured nozzles. It is found that they fail to predict the separation location correctly, exhibiting sensitivity to the range of NPRs and to the type of nozzle. Depending on NPR, the SST model either under-predicts or over-predicts Free Shock Separation (FSS). Moreover, it also fails to capture Restricted Shock Separation (RSS). With compressibility correction, it under-predicts separation at all NPRs to a greater extent. Even though RSS is captured by using compressibility correction, the transition from FSS to RSS is over-predicted [2]. Early efforts by few researchers to improve predictions of nozzle flow separation by realizability corrections to turbulence models have not been successful, especially in terms of capturing both the separation types. Therefore, causes of turbulence modeling failure in predicting nozzle flow separation correctly were further investigated. It is learnt that limiting of the shear stress inside boundary layer, due to Bradshaw’s assumption, and over-prediction of jet spreading rate are the causes of SST model’s failure in predicting nozzle flow separation correctly. Based on this physical reasoning, values of the a 1 parameter and the two diffusion coefficients σk,2 and σω,2 were empirically modified to match the predicted wall pressure distributions with experimental data of the DLR-PAR and the VAC-S6-short nozzles. The results confirm that accurate prediction of flow separation in rocket nozzles indeed depends on the correct prediction of spreading rate of the supersonic separation-jet. It is demonstrated that accurate RANS simulation of flow separation in rocket nozzles over a wide range of NPRs is feasible by modified values of the diffusion coefficients in turbulence model.

Methane-based rocket propellant is gaining traction as a green technology with advantages in sustainability, cost-effectiveness, and performance. However, under high temperatures found in rocket nozzle cooling channels, methane can undergo thermal decomposition known as methane pyrolysis, resulting in the generation of hydrogen and solid carbon. This poses challenges to rocket engine performance and can eventually cause engine failure. Understanding and predicting the composition of evolved gases in rocket engine processes is therefore crucial. This thesis focuses on quantifying the production of hydrogen in the exhaust stream. To achieve this objective, a correlational measurement method utilizing sensors was developed and experimentally investigated. This approach involved the detailed mapping of sensor responses to variations in gas composition, temperature, and pressure, which were compared and validated against theoretical data derived from REFPROP; a widely used software tool for calculating gas properties. The sensors employed in this study enabled direct measurements of the speed of sound (SOS) and thermal conductivity (TCD) of the gas. The SOS measurements exhibited strong agreement with theoretical predictions in response to changes in hydrogen content. In contrast, the TCD measurements showed lower sensitivity to hydrogen. It was observed that temperature exhibited a substantial influence on both SOS and TCD compared to pressure. However, the implementation of experimental and theoretical correction coefficients effectively compensated for these effects. The resulting calibration curves demonstrated an absolute deviation of 0.2-0.3%vol in hydrogen concentration, which demonstrates the effectiveness of the developed method of quantifying hydrogen in gas mixtures. Lastly, the occurrence of methane pyrolysis was tested and confirmed. / Metan-baserat raketbränsle är en attraktiv grön teknologi med fördelar inom hållbarhet, kostnadseffektivitet och prestanda. Dock kan metan vid höga temperaturer funna i kylningskanalerna av raketmunnstycken undergå termisk sönderfallning via en process som kallas för metanpyrolys, vilket resulterar i produktionen av vätgas och fast kol. Detta medför utmaningar för prestandan av raketmotorn och kan i slutändan förstöra motorn. Det är därför mycket viktigt att kunna förstå och förutsäga sammansättningen av de gaser som bildas i processerna i raketmotorer. Detta examensarbete fokuserar på att kvantifiera produktionen av vätgas i avgasströmmen. För att uppnå detta mål utvecklades och experimentellt undersöktes en korrelationsmätmetod som använder sensorer. Detta tillvägagångssätt innebar en detaljerad kartläggning av sensorernas svar på variationer i gassammansättning, temperatur och tryck, som sedan jämfördes och validerades mot teoretiska data från REFPROP; ett välkänt programverktyg för beräkningen av gasegenskaper. De sensorer som användes i denna studie möjliggjorde direkta mätningar av ljudhastigheten (SOS) och värmeledningsförmågan (TCD) hos gasen. SOS-mätningarna visade en stark överensstämmelse med teoretiska förutsägelser som svar på förändringar i vätgasinnehållet. TCD-mätningarna visade däremot lägre känslighet för väte. Det observerades att temperaturen hade en betydande inverkan på både SOS och TCD jämfört med trycket. Implementeringen av experimentella och teoretiska korrigeringskoefficienter kompenserade dock effektivt för dessa effekter. De resulterande kalibreringskurvorna visade en absolut avvikelse på 0.2-0.3%vol i vätgaskoncentration, vilket betonar effektiviteten hos den utvecklade metoden för att kvantifiera väte i gasblandningar. Slutligen testades och bekräftades förekomsten av metanpyrolys.