An approach for designing a high-enthalpy wind tunnel driven by exothermic chemical reactions was developed. Nozzle contours were designed using CONTUR, a program implementing the method of characteristics, to design nozzle contours at various flow conditions. A reacting mixture including nitrous oxide has been identified as the best candidate for providing clean air at high temperatures. The nitrous oxide has a few performance factors that were considered, specifically the combustion of the gas. Initial CFD simulations were performed on the nozzle and test region to validate flow characteristics and possible issues. Initial results show a fairly uniform exit velocity and ability to perform testing. In a second phase of the work, two generic, high-speed missile configurations were explored using numerical simulation. The mean flow was computed on both geometries at 0 and 45 roll and 0, 1, and 10 angle of attack. The computations identified complex flow structures, including three-dimensional shock/boundary-layer interactions, that varied considerably with angle of attack. <br>

<div>Shock-wave/turbulent boundary-layer interactions (SWTBLIs) are characterized by low-frequency unsteadiness, amplified aerothermal loads, and a complex three-dimensional flowfield. Presence of a broad range of length and time-scales associated with compressible turbulence generates additional gasdynamic features that interact with different parts of the flowfield via feedback mechanisms. Determining the physics of such flows is of practical importance as they occur frequently in different components of a supersonic/hypersonic aircraft such as inlets operating in both on- and off-design conditions, exhaust nozzles, and control surfaces. SWTBLIs can cause massive flow separation which may trigger unstart by choking the flow in an inlet. On control surfaces, fatigue loading caused by low-frequency shock unsteadiness, coupled with high skin-friction and heat transfer at the surface, can result in failure of the structure.</div><div><br></div><div>The objective of this study is twofold. The first aspect involves examining the causes of unsteadiness in SWTBLIs associated with two geometries – a backward facing step flow reattaching on to a ramp, and a highly confined duct flow. Signal processing and statistical techniques are performed on the results obtained from Delayed Detached-Eddy Simulations (DDES) and Implicit Large-Eddy Simulations (ILES). Dynamic Mode Decomposition (DMD) is used as a complement to this analysis, by obtaining a low-dimensional approximation of the flowfield and associating a discrete frequency value to individual modes. </div><div><br></div><div>In case of the backward facing step, Fourier analysis of wall-pressure data brought out several energy dominant frequency bands such as separation bubble breathing, oscillations of the reattachment shock, shear-layer flapping, and shedding of vortices from the recirculation zone. The spectra of reattachment shock motion suggested a broadband nature of the oscillations, wherein separation bubble breathing affected the low-frequency motion and shear-layer flapping, and vortex shedding correlated well at higher frequencies. A similar exercise was carried out on the highly confined duct flow which featured separation on the floor and sidewalls. In addition to the low-frequency shock motions, the entire interaction exhibited a cohesive back-and-forth in the streamwise direction as well as a left-right motion along the span. Mode reconstruction using DMD was used in this case to recover complex secondary flows induced by the presence of sidewalls.</div><div><br></div><div>For the final aspect of this study, a flow-control actuator was computationally modeled as a sinusoidally varying body-force function. Effects of high-frequency forcing at F⁺ =1.6 on the flowfield corresponding to a backward facing step flow reattaching on to a ramp were examined. Conditionally averaged profile of streamwise velocity fluctuations, based on reattachment shock position, was used for the formulation of spatial distribution of the actuator. The forcing did not change the mean and RMS profiles significantly, but affected the unsteadiness of the interaction significantly. The effects of forcing were localized to the recirculation zone and did not affect the evolution of the shear-layer. The acoustic disturbances propagating through the freestream and recirculation zone drove the motion of the reattachment shock, and did not alter the low-frequency dynamics of the interaction.</div>

This research focuses on measurements of a convective shear-layer instability seen naturally in quiet hypersonic flow. Experiments were carried out in the Boeing/AFOSR Mach 6 Quiet Tunnel (BAM6QT) at Purdue University. The BAM6QT provides low-disturbance hypersonic flow with freestream noise levels similar to what would be experienced by a flight vehicle. To obtain high-speed, off-the-surface measurements of the instability, a modified focused laser differential interferometer (FLDI) was first designed to work with the contoured Plexiglas windows available in the tunnel.<div><br>A cone-cylinder-flare geometry was then selected to study the instabilities related to an axisymmetric separation bubble at Mach 6. The sharp cone had a 5-degree half-angle, while flare angles of 10 degrees and 3.5 degrees were tested to compare axisymmetric compression with and without separation, respectively. Under quiet flow, laminar separation and reattachment was confirmed by schlieren and surface pressure-fluctuation measurements. Coherent traveling waves were observed. These were attributed to both the second-mode instability, as well as a shear-generated instability from the separation bubble. The symmetry of the bubble was found to be highly sensitive to angle of attack. Additionally, by introducing controlled disturbances on the cone upstream of the separation, larger-amplitude shear-generated waves were measured while the second-mode amplitudes remained unchanged. Therefore, the shear-generated waves were amplified moving through the shear layer, while the second mode remained neutrally stable. These appear to be the first measurements of traveling waves that are generated in the shear layer of a separation bubble in hypersonic flow. <br></div>

<p>Air intakes are a fundamental part of all high speed airbreathing propulsion concepts. The main purpose of an intake is to capture and compress freestream air for the engine. At hypersonic speeds, the intake’s surface and shock structure effectively slow the airflow through ram-air compression. In supersonic-combustion ramjets, the captured airflow remains supersonic and generates complicated shock structures. The design of these systems require careful evaluation of proposed operating conditions and relevant aerodynamic phenomena. The physics of these systems, such as the intake’s operability range, mass capture efficiency, back-pressure resiliency, and intake unstart margins are all open areas of research. </p> <p><br></p> <p>A high speed intake, dubbed the Indiana Intake Testbed, was developed for experimentation within the Boeing-AFOSR Mach 6 Quiet Tunnel at Purdue University. This inward-turning, mixed compression intake was developed from osculating axisymmetric theory and uses a streamtracing routine to create a shape-transitioned geometry. To account for boundary layer growth, a viscous correction was implemented on the intake’s compression surfaces. This comprehensive independent design code was pursued to generate an unrestricted geometry that satisfies academic inquiry into fluid dynamic interactions relevant to intakes. Additionally, the design code contains built-in analysis tools that are compared against CFD calculations and experimental data. </p> <p><br></p> <p>Two blockage models were constructed and outfitted with Kulite pressure transducers to detect possible intake start and unstart effects. Due to an error in the design code, the preliminary blockage models’ lower surfaces were oversized. The two intake models were tested over a freestream Reynolds number sweep, under noisy and quiet flow, at one non-zero angle of attack, and at a singular back-pressure condition. Back-pressure effects acted to unstart the intake and provide a comparison between forced-unstart and started states. The experimental campaign cataloged both tunnel starting and inlet starting conditions, which informed the design of the finalized model. The finalized model is presented herein. Future experiments to study isolator shock-trains, shock-wave boundary layer interactions, and possible instances of boundary layer transition on the intake’s compression surface are planned. </p>

<p> The Virtual Super-resolution Optics with Reconfigurable Swarms (VISORS) mission  aims to produce high-resolution images of solar release sites in the solar corona using a  distributed telescope. The collected data will be used to investigate the existence of underlying  energy release mechanisms. The VISORS telescope is composed of two spacecraft flying in a  formation configuration. The optics spacecraft (OSC) hosts the optic system, while the detector  spacecraft (DSC) is located behind the OSC in alignment with the Sun and houses a detector.  The two modes of operation for the CubeSats are Science Operations Mode and Standby Mode.  In Science Operations Mode, the two spacecraft are at a close distance which may make the plume impingement an issue. The cold gas thruster propulsion systems in both the OSC and  DSC use R-236fa (HFC) refrigerant. The plume from the system is modeled using SPARTA  Direct Simulation Monte Carlo (DSMC) Simulator while the refrigerant itself is modeled using  an equivalent particle that closely matches viscosity and specific heat. This work aims to  investigate plume propagation for two different flow inputs. The DSMC simulations are  performed with the input parameters acquired using the isentropic relations and CFD simulations  of the 2D axisymmetric nozzle flow. Additionally, the DSMC results are compared to the  Boynton-Simons, Roberts-South, and Gerasimov analytical plume models. </p>

<p dir="ltr">A fast model of a Rotating Detonation Combustor (RDC) is developed based on the Method of Characteristics (MOC). The model provides a CFD-like solution of an unwrapped 2D RDC flow field in under 10 seconds with similar fidelity as 2D Reacting URANS simulations. Parametric studies are conducted using the simplified model, and the trends are analyzed to gain insight into the underlying physics of rotating detonation combustors. A methodology to assess the performance of operation with multiple waves is presented. The main effect of increasing waves is found to be the increase in the exit Mach number of the combustion chamber. The design process of a mixer component is also presented. The mixer lies downstream of a channel-cooled RDC with subsonic exit and upstream of a Rolls-Royce M250 helicopter engine in open-loop configuration. The mixer dilutes the RDC exhaust with approximately 250% air to condition the flow for the M250 turbine at steady state operation, while also acting as an isolator with a choked throat to prevent back propagation of pressure waves. The mixer aerodynamic design was completed using 2D axisymmetric RANS simulations, and the mechanical design was evaluated using Ansys Mechanical FEA and was found to be able to survive the high thermal stresses present both during the transient heating and steady state operating condition.</p>

<p>In rotating detonation engines the turbine inlet conditions may be transonic with unprecedented unsteady fluctuations, very different from those in conventional high-pressure turbines. To ensure an acceptable engine performance, the turbine passages must be unchoked at subsonic and started at supersonic conditions. Additionally, to maximize the aerodynamic performance potential, ad-hoc designs are required, suited for the oscillations in Mach number and flow angle. This manuscript focuses on designing and characterizing diffusive turbine vanes that can operate downstream of a transonic rotating detonation combustor.  </p> <p><br></p> <p>First, the phenomenon of unstarting is presented, concentrating on the effect of pressure loss on the accurate prediction of the starting limit. Afterward, a multi-objective optimization with steady Reynolds Averaged Navier Stokes simulations, including the endwall and 3D vane design, is performed. The results are discussed, highlighting the impact of the throat-to-inlet area ratio on the pressure loss and the geometric features of the top-performing designs. Compared to previous  research on stator passages with contoured endwalls, considerable reductions in pressure loss and stator-induced rotor forcing are obtained, with an extended operating range and preserving high turning.  </p> <p><br></p> <p>Subsequently, the influence of the inlet boundary layer thickness on the vane performance is evaluated, inducing remarkable increases in pressure loss and downstream pressure distortion. Employing an optimization with a thicker inlet boundary layer, specific endwall design recommendations are found, providing a notable improvement in both objective functions. The impact of the geometry variations on flow detachment is assessed as well.</p> <p><br></p> <p>Finally, the impact of the inlet flow angle on the vane design is studied through a multi-point, multi-objective optimization with different inlet angles. The effect of incidence on the flow field and vane performance is evaluated first. Then, by comparing the optimized geometries with those optimized for axial inflows, several design guidelines are identified </p>

<p>Purdue University is facing an increasing undergraduate student enrollment every year. Laboratory courses in fluid mechanics at the School of Aeronautics and Astronautics and the School of Mechanical Engineering are facing challenges due to increased enrollments. On the other hand, the internet and computer technology have made education more conveniently delivered in recent years. The onset of these technologies has made way for innovative forms of teaching. One such application is virtual laboratories. This document will describe the challenges in the current method of teaching and learning fluid mechanics laboratory courses, explain how a virtual lab is a potential solution to supplement the current learning methods, and discuss its development. This document will discuss the virtual fluid mechanics laboratory development from the fluid mechanics perspective- generating flow data and designing realistic measurement probes. The use of CFD to generate flow data sets, along with their post-processing for virtual labs, will be discussed. Furthermore, simulation results for flow around cylinders and through pipes will be presented. And finally, design ideas conceptualized for developing a virtual pitot-static probe and a virtual hot-wire anemometer will also be presented. Finally, this document summarizes the work done till now and presents conclusions on what has been achieved along with recommendations that could be completed in the future.</p>

<p dir="ltr">On many hypersonic vehicles, gaps are present on the outer surface of the vehicle and the interaction of the hypersonic freestream flow over these gaps can cause significant heat transfer to the vehicle. The project described in this thesis analyzed selected hypersonic gap problems and attempted to offer solutions to combat the heat transfer occurring in the gap. The first section of this thesis is a parametric study to understand the changes to the heat transfer and flow that modifications to the gap geometry can make. The second section is a comparison of the computational model to experimental data. The results of the studies show that adding a simple fillet or chamfer to the downstream step of the gap can reduce the maximum heat flux by over 90%. These results can be used to reduce the heat transfer caused by flow impingement in the gaps of hypersonic vehicles with a simple modification of the geometry and is consistent with the findings of other work in gap heating.</p>

<p dir="ltr">Increased scrutiny from customers and regulators to design aeroengines that are more efficient and environmentally friendly has pushed the need to investigate new engine architectures, manufacturing techniques, and computational fluid dynamic methods. This has led to the development of the CSTAR Gen. 2.5 centrifugal compressor, which uses an additively manufactured diffusion system and investigates the aerodynamic performance of an axi-centrifugal aeroengine. Additionally, an isolated unsteady diffuser computational model was previously developed that seeks to significantly reduce the computational cost of unsteady CFD in the diffuser.</p><p dir="ltr">The research presented in this paper is part of an ongoing attempt to utilize the capabilities of isolated unsteady diffuser modeling and rapid prototyping enabled through additive manufacturing in CSTAR Gen. 2.5 to develop a design framework that allows for quick computational and experimental analysis of diffusion systems in centrifugal compressors. Specifically, the scope of isolated unsteady diffuser modeling, which was previously only implemented in CSTAR Gen. 1 and at a single loading condition, is expanded by analyzing computational instabilities when applying the methodology to CSTAR Gen. 2.5 and analyzing results from four loading conditions (high loading, design point, low loading, and near choke) along a speedline.</p><p dir="ltr">Computational instabilities in the CSTAR Gen. 2.5 isolated diffuser models were determined to be caused by the decreased vaneless space compared to Gen. 1, which led to less “mixed” flow at the impeller outlet and a stronger diffuser potential field affecting the inlet profile. A boundary profile correction approach was developed which slightly increased very low total pressure near the diffuser shroud and negative radial velocity regions near the shroud and pitchwise locations of the diffuser vane leading edges while minimizing the overall affected area. The correction was successfully validated using 3D flow structure and minimum, average, and maximum total pressure, absolute velocity magnitude, and pressure comparisons at the diffuser inlet between an isolated and full-stage model.</p><p dir="ltr">Prediction capabilities of 3D flow structures and 1D performance parameters by isolated unsteady diffuser models were validated with results from full-stage unsteady models at each loading condition. The analysis indicated consistent performance by the isolated unsteady diffuser model at all loading conditions. An overall agreement in 3D flow structures was found, and trends in the full-stage unsteady models along the speedline were tracked well by the isolated unsteady model. At all loading conditions, there was a consistent over-representation of the separation region along the diffuser vane pressure side in the diffuser passage, overprediction of total pressure magnitude at the core of the flow at the diffuser outlet, and over- or underprediction of total pressure loss and static pressure recovery respectively. The similarity in the results between full-stage and isolated unsteady models, tracking of trends along the speedline, and consistent differences in 3D flow structure predictions and 1D performance parameters validates the isolated unsteady diffuser methodology for use at loading conditions from surge to choke.</p>