To further investigate the formation and behavior of streamwise heat streaks in hypersonic flow, a simplified highly-swept model with a second-order continuous leading edge was tested in the Actively Controlled Expansion (ACE) hypersonic tunnel at Texas A&M University. The study aimed to examine the effects of shock curvature and the influence of second-order leading edge continuity on heat streak behavior at Mach 5.7 across a variety of pitch angles and Reynolds numbers. Diagnostic techniques for analysis included high-speed schlieren, infrared thermography, and surface-mounted pressure transducers. The results showed that heat streaks developed along the test article's surface, with their intensity increasing with both Reynolds number and angle of attack. The streaks were less intense and appeared further downstream compared to those observed on geometries with second-order discontinuous leading edges under identical conditions. In addition, their spanwise location shifted inward with increasing angle of attack. Surface pressure spectra revealed broadband disturbances from 15 to 50 kHz at various positions along the model, with characteristic frequencies remaining constant but intensities scaling with flow conditions. These findings suggest that, while second-order continuity of the leading edge may reduce initial shock curvature concentrations, it does not solely eliminate the generation of vortical structures responsible for localized heating in laminar hypersonic flows. Spanwise pressure gradients and the turning radius of the corner, regardless of the continuity of the leading edge, play a key role in the formation of heat streaks.

The Mach 6 flow over a blunt wedge is analyzed, with the aim to uncover the origin of a hot streak that forms on top of the wedge. Vortex identification criteria link the hot streak to a streamwise vortex, that produces a downwash at the location of the hot streak. The hot streak is therefore the likely consequence of the thus-enhanced frictional heating. One of the vortex identification criteria indicates that the shock is involved in the creation of the streamwise vortex, another considered criterion does not, however. A more advanced vortex analysis approach should be used to unequivocally determine the role the shock plays in the formation of the streamwise vortex on top of the wedge. Besides the shock, a strong spanwise pressure gradient along the surface of the wedge may be responsible for the creation of the streamwise vortex. In fact, an 'inboard' and 'outboard' streamwise vortex are identified, dwelling on opposite sides of the pressure minimum in the spanwise direction.

A simplified, swept-wedge model was tested in the Actively Controlled Expansion (ACE) Hypersonic Tunnel at Texas A&M University over a range of freestream conditions to isolate the origin of heat streaks emanating from a junction between the leading edge and the swept-sides of the model. High-speed schlieren imagery, infrared thermography, and surface-mounted transducers were used to characterize the heat streaks at several angles of attack and Reynolds numbers. A complementary computational study informed the placement of instrumentation relative to predicted streamwise streak locations. Schlieren imagery detailed the shape and structure of the bow shock and boundary layer, while infrared thermography showed two prominent heat streaks emanating from the junction between the leading edge and swept-sides of the model. Streak intensity was proportional to increases in angle of attack and Reynolds number. Power spectra for the surface pressure sensors observed broadband frequency responses in the range of 10 - 50 kHz under the streak and inboard toward the centerline. Spectra for each transducer showed that characteristic frequencies were invariant with angle of attack, while their intensity was proportionally related.

High-fidelity computational fluid dynamics of hypersonic flow over a swept wedge model were performed to investigate the origin mechanisms of heat streaks emanating from the junction between the leading edge and swept sides of a wedge geometry. The computational campaign examined the impact of Reynolds number, leading edge curvature, and angle of attack on streak structure. Additionally, a study of flow-field sensitivity to different numerical methods and grid resolutions was completed. CFD simulations were also used to inform test article design and placement of instrumentation for a complementary experimental campaign performed in the Actively Controlled Expansion (ACE) wind tunnel at Texas A&M University. Heat streak structures appear highly sensitive to Reynolds number, leading edge curvature, and angle of attack. Possible streak generation mechanisms include a combination of vorticity driven by shock curvature and non-continuous surface curvatures at the leading edge junction. Future work will seek to isolate these mechanisms.

Hypersonic boost-glide vehicles are designed to fly long distances in the upper atmosphere. They are reported to have the potential to evade ballistic missile early warning systems and to maneuver as they fly toward their target. A recent analysis by Tracy and Wright in Science & Global Security claimed to show that typical boost-glide vehicles produce significant infrared signatures that would be readily detectable with existing U.S. satellites and therefore questioned the potential advantages of hypersonic weapons over existing missiles. The prior analysis is revisited and several inconsistencies in the underlying assumptions are described. A detailed computational fluid dynamics analysis predicts typical infrared signatures to be significantly lower than those predicted by Tracy and Wright. As a result, these signatures would fall below the detection threshold of legacy U.S. Defense Support Program satellites but remain detectable by the more modern sensors from the Space-Based Infrared System. There are two significant issues with the prior analysis: an incorrect aerodynamic angle of attack was used, and the turbulent heat transfer rate correlation used to predict the surface temperature is inaccurate at the conditions studied.

Covers advancements in spacecraft and tactical and strategic missile systems, including subsystem design and application, mission design and analysis, materials and structures, developments in space sciences, space processing and manufacturing, space operations, and applications of space technologies to other fields.

This study explores the structure of liquid/gas coaxial jets under forced and unforced conditions. The forcing is in the form of a transverse acoustic resonance within the confined space where the mixing occurs. The studied flows are relevant to combustion instabilities which involve an interaction between acoustic waves and reactant mixing. A variety of local and global signal processing methods were applied to digital flow visualization data to identify spatial and temporal features. The unforced case is in particular chaotic and influenced by a broad range of spatial and temporal phenomena. Proper orthogonal decomposition (POD) was able to extract flapping and convecting features, and spectral content of these behaviors is presented. The forced case results in organized structures that emerge above the background turbulence, including harmonics of the forcing frequency and nonlinear interactions between specific frequencies. The dynamic mode decomposition (DMD) performs the best in the forced case, clearly isolating all of these features. Wavelet analysis showed that forcing tended to reorganize energy from longer to shorter time scales. Bicoherence analysis of the data showed that the forcing causes a much different energy exchange in the outer and inner shear layers. The outer-to-inner jet coupling during forced conditions appears to be limited to an axial extent of about one to three inner jet diameters downstream of the jet exit. The recirculation zone between the inner and outer jet, extending about one inner jet diameter downstream, appears to disrupt the influence of forcing on the inner jet.

The turbulent spot propagation process in boundary layer flows of air, nitrogen, carbon dioxide, and air/carbon dioxide mixtures in thermochemical nonequilibrium at high enthalpy is investigated. Experiments are performed in a hypervelocity reflected shock tunnel with a 5-degree half-angle axisymmetric cone instrumented with flush-mounted fast-response coaxial thermocouples. Time-resolved and spatially demarcated heat transfer traces are used to track the propagation of turbulent bursts within the mean flow, and convection rates at approximately 91, 74, and 63% of the boundary layer edge velocity, respectively, are observed for the leading edge, peak, and trailing edge of the spots. A simple model constructed with these spot propagation parameters is used to infer spot generation rates from observed transition onset to completion distance. Spot generation rates in air and nitrogen are estimated to be approximately twice the spot generation rates in air/carbon dioxide mixtures.

How much new science will it take to design a vehicle that can routinely fly at many times the speed of sound?

No AccessTechnical NoteEffects of Shock-Tube Cleanliness on Hypersonic Boundary Layer Transition at High EnthalpyJoseph S. Jewell, Nicholaus J. Parziale, Ivett A. Leyva and Joseph E. ShepherdJoseph S. JewellU.S. Air Force Research Laboratory, Wright–Patterson Air Force Base, Ohio 45433, Nicholaus J. ParzialeStevens Institute of Technology, Hoboken, New Jersey 07030, Ivett A. LeyvaU.S. Air Force Office of Scientific Research, Arlington Air Force Base, Virginia 22203 and Joseph E. ShepherdCalifornia Institute of Technology, Pasadena, California 91125Published Online:16 Sep 2016https://doi.org/10.2514/1.J054897SectionsRead Now ToolsAdd to favoritesDownload citationTrack citations ShareShare onFacebookTwitterLinked InRedditEmail About References [1] Bushnell D., "Notes on Initial Disturbance Fields for the Transition Problem," Instability and Transition, edited by Hussaini M. and Voigt R., ICASE/NASA LaRC Series, Springer, New York, 1990, pp. 217–232. 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: A high-pressure nano-ignition torch has been developed which takes advantage of the photoignition properties of single wall carbon nanotubes (SWNTs). The initiation of combustion in a cryogenic O2-H2 coaxial injector at about 32 atm (470 psi) with O2 temperatures of about 130 K was achieved with a 200 mg torch containing no more than 70 mg of SWNT-based fuel. Our investigation includes the effects of chamber pressure, the presence of different solid oxidizers such as boronpotassium nitrate (BKNO3), potassium permanganate (KMnO4)and ammonium perchlorate (AP) as well as solid fuels, such as aluminum nanoparticles and solid propellant, on the functionality of the nano-ignition torch. We have shown that by mixing SWNT with other nanoparticles and powdered materials, the ignition parameters such as burn temperature, burn duration and the ignition byproducts can be tailored to meet different ignition requirements. It is anticipated that photoignition provides a suitable method for ignition of systems that require the start of combustion at a high pressure up to about 135 atm (2000 psi) as well as ignition of certain thrusters and liquid rocket engines that require an extremely small and light weight torch igniter. The disclosed ignition method is based on robust off-the-shelf technology and it is scalable to applications in very small as well as large combustion chambers such as gas turbines, gas generators, liquid rocket engines and possibly multi grain solid rocket motors.

Abstract : A high-pressure photoignition torch has been developed which takes advantage of the photoignition properties of single wall carbon nanotubes (SWNTs). The goal was to initiate combustion in a cryogenic O2-H2 coaxial injector at about 35 atm (520 psi) at O2 temperature of about 130 K with SWNT-based solid fuel mixtures. Our investigation includes the effects of chamber pressure, the presence of different solid oxidizers such as BKNO3 and KMnO4, as well as solid fuels and solid propellants, on the functionality of the photoignition torch. We have shown that the ignition parameters such as burn temperature, burn duration and the ignition byproducts can be tailored to meet different ignition requirements. It is anticipated that photoignition provides a suitable method for ignition of systems that require the start of combustion at a high pressure up to about 135 atm (2000 psi) as well as ignition of certain thrusters and liquid rocket engines that require an extremely small and light weight torch igniter. This ignition method can be applied to large combustion chambers such as gas turbines, gas generators, liquid rocket engines and possibly multi grain solid rocket motors.

The Caltech T5 reflected shock tunnel is used to produce hypervelocity flow over a range of velocities and pressures by varying the test gas and operating parameters of reservoir enthalpy (hres) and reservoir pressure (Pres). One area of research in T5 is the measurement of boundary layer behavior and transition from laminar to turbulent flow on a smooth 5-degree half-angle cone [3, 1, 11].

A series of slender-body hypervelocity boundary-layer instability and transition experiments were performed in the Caltech T5 Reflected-Shock Tunnel. During this campaign, it became clear that the condition of the T5 shock tube would significantly affect the consistency of the instability and transition measurements; a regimen of cleaning was iterated on until satisfactory repeatability was achieved. In this work, a description of the cleaning regimen is given. Additionally, boundary-layer instability measurements and a statistical analysis of the boundary-layer transition scatter are presented for experiments before and after cleaning regimen implementation.

: An experimental study has been conducted to explore the coupling between a coaxial gaseous hydrogen / liquid oxygen jet flame and transverse acoustic perturbations. A variety of chamber conditions including acoustic frequency, amplitude, and the location of the pressure node / antinode with respect to the flame were examined. The flame response was documented using high-speed imaging including backlit visualization and unfiltered chemiluminescence. Dynamic mode decomposition was used to isolate the spatial structure of the flame response at the forcing frequency. The results indicate that the flame response to forcing is qualitatively similar to previous results of nonreacting coaxial jet flows; the pressure node forcing appears to generate in-plane flapping of the flame while pressure antinode forcing induces a helical structure in the flame.

: An experimental study has been conducted at the Air Force Research Laboratory at Edwards Air Force Base to explore the receptivity of cryogenic coaxial jet flows to transverse acoustic disturbances. The shear coaxial jet flow employed liquid nitrogen in the inner jet and cooled helium in the outer annular jet to represent the nominal fluid dynamical conditions of an oxygen/hydrogen liquid rocket engine injector. The injector flow is submerged in a chamber that experiences a monotonic transverse acoustic resonance characteristic of a rocket chamber in the presence of combustion instability. The coaxial jet is exposed to a variety of acoustic conditions including different frequencies, amplitudes, and locations within the resonant mode shape. High-speed back-lit images were captured to record the behavior of the natural (unforced) and forced coaxial jets. Proper orthogonal decomposition and spectral analysis were used to extract natural and forced modes. Convective modes are extracted, and a new Strouhal number is used to characterize the dominant natural convective mode that is analogous to the preferred mode in free jets. The threshold of receptivity was found for a number of different injector flows and acoustic forcing conditions. The results indicate that the dimensionless frequency plays an important role, and there exists a finite forcing amplitude at which the threshold of receptivity occurs. The receptivity threshold and post receptivity response provides useful insight on the suitability of a given injector design for specific rocket combustion chamber conditions.

Abstract : A novel method to delay transition in hypervelocity flows by using non-equilibrium effects has been studied under Grant 13RQ14COR. The first molecule studied was carbon dioxide (CO2). The motivation arises from experimental and numerical data showing that when pure CO2 is in vibrational and chemical non-equilibrium, these relaxation processes absorb energy from acoustic disturbances whose growth in the boundary layer is responsible for transition through the 2nd or Mack mode in certain hypervelocity configurations. By absorbing energy at the same frequencies of the acoustic disturbances, nonequilibrium CO2 delays transition in hypervelocity flows. Before this program, no effort had been made to extend these results to CO2 injection into base air boundary layers.

: Combustion instability in liquid rocket engines can have severe consequences including degraded performance, accelerated component wear, and potentially catastrophic failure. High-frequency instabilities, which are generally the most harmful in liquid rocket engines, can be driven by interactions between disturbances associated with transverse acoustic resonances and the combustion process. The combustion response to acoustic perturbation is a critical component of the instability mechanism, and is in general not well understood. The current paper describes an experimental facility at the Air Force Research Laboratory (AFRL) at Edwards Air Force Base that is intended to investigate the coupling between transverse acoustic resonances and single/multiple liquid rocket engine injector flames. Critical aspects of the facility will be described, including the capability to operate at supercritical pressures that are relevant to high-performance liquid rocket engines, accurately-controlled and cryogenically-conditioned propellants, and optical access to facilitate the use of advanced diagnostics. The transverse acoustic resonance is induced through the use of carefully-controlled piezo-sirens, allowing monochromatic excitation across a range of amplitudes at a number of discrete frequencies. The location of the flame within the acoustic resonance mode shape can also be varied through relative phase control of the two acoustic sources. The operating space of the facility, for oxygen and hydrogen operation, will be described. Preliminary non-reacting and reacting data will also be presented to demonstrate the quality of operation of this facility. It is anticipated that future results generated using this facility will provide both fundamental insight into the acoustic-flame interactions as well as provide a database useful for validating combustion instability models.