AIAA Journal Research Articles

Improved Design Method for Hypersonic Intakes Using Pressure-Corrected Osculating Axisymmetric Flows Method

The efficient design of air intake is crucial for high-speed airbreathing propulsion systems. This paper presents a novel design method that integrates the recently introduced internal conical flow M (ICFM) basic flowfield (Musa et al., “New Parent Flowfield for Streamline-Traced Intakes,” AIAA Journal, Vol. 61, No. 7, 2023, pp. 2906–2921) with the pressure-corrected osculating axisymmetric flows and streamline tracing techniques. The new design method takes into account the effects of lateral pressure gradients by incorporating pressure corrections into the original design method using crossflow velocity information from the ICFM flowfield. Additionally, new entrance and throat profiles are introduced for designing two hypersonic internal waverider intakes with shape transition. One intake is designed using the new method, and the other using the original method. Viscous simulations have been performed at design (i.e., Mach 6.0) and off-design (i.e., Mach 4.0 and 3.5) conditions. The results indicate that the pressure-corrected intake exhibits superior performance in terms of total pressure recovery and reduced flow distortion. This reveals that the new design method can achieve a smooth shape and efficient aerodynamic transitions between the intake entrance and throat. The new entrance and throat profiles realized better performance than the typical ones. The startability analysis reveals that the Van Wie curve controls the considered intakes. Thus, combining the pressure-corrected osculating axisymmetric flows with the ICFM flowfield enhances the performance of hypersonic intakes, making this approach a promising avenue for future hypersonic vehicles.

Direct Numerical Simulation of High-Enthalpy Turbulent Boundary-Layer Flow with Light Gas Injections

Two light gases (He and H2) are, respectively, introduced upstream of a high-enthalpy turbulent flat-plate flow with a boundary-layer edge Mach number of Maδ=2.25 and temperature of Tδ=1800 K. The flow condition refers to the after-shock wave flow on a blunt-body hypersonic vehicle (Duan and Martín, AIAA Journal, Vol. 49, No. 1, 2011, pp. 172–184). Direct numerical simulation results show that the injection of these light gases has little effect on the mean velocity profiles but significantly reduces the near-wall density and skin friction. The separation and fragmentation of near-wall vortical structures are restrained, and the Reynolds shear stress decreases. The injection of inert He inhibits the dissociation reaction of O2 and weakens the chemical nonequilibrium effect, resulting in enhanced mean and fluctuating temperatures. The injection of active H2 promotes the reaction between H2 and O2, which increases the mean temperature but inhibits its fluctuation. After decomposing the mean skin friction into physics-informed contributions, both injections largely reduce the turbulent kinetic energy production term Cf,T and the spatial growth term of the flow, Cf,G, through lowering the near-wall density and reducing vortices.

Convolutional neural networks for tracking coalescing waveforms in supersonic jet noise

High-speed schlieren images of the field surrounding a supersonic jet provide a rich database for testing convolutional neural network (CNN) classification methods due to the presence of many waveform structures of interest. One such structure is waveform coalescence, where waves intersect at small angles, which can lead to increased steepening and nonlinear distortion in the jet near field [Willis et al., AIAA Journal (2023)]. Although waves exhibiting coalescence behavior have been identified in narrow field-of-view (FOV) schlieren images, new methods are needed for large FOV images to improve computational efficiency. This presentation will explore methods to track and classify waves observed in large FOV schlieren images as coalescing or noncoalescing using CNNs. Using transfer learning, pretrained networks can be retrained for this problem, with training data obtained from narrow FOV schlieren or from two-dimensional simulated waveforms using the Khokhlov–Zabolotskaya–Kuznetzov (KZK) equation. The impact of model hyperparameters, choice of pretrained network, image scaling, and other factors on the results of waveform classification by the CNN will be explored. [WAW is supported by the ARL:UT Chester M. McKinney Graduate Fellowship in Acoustics.]

Sensitivity Analysis of Direct Numerical Simulation of a Spatially Developing Turbulent Mixing Layer to the Domain Dimensions

Abstract Understanding spatial development of a turbulent mixing layer is essential for many engineering applications. However, the flow development is difficult to replicate in physical or numerical experiments. For this reason, the most attractive method for the mixing layer analysis is the direct numerical simulation (DNS), with the most control over the simulation inputs and free from modeling assumptions. On the other hand, the DNS cost often prevents conducting the sensitivity analysis of the simulation results to variations in the numerical procedure and thus, separating numerical and physical effects. In this paper, effects of the computational domain dimensions on statistics collected from DNS of a spatially developing incompressible turbulent mixing layer are analyzed with the focus on determining the domain dimensions suitable for studying the flow asymptotic state. In the simulations, the mixing layer develops between two coflowing laminar boundary layers formed on two sides of a sharp-ended splitter plate of a finite thickness with characteristics close to those of the untripped boundary layers in the experiments by Bell and Mehta, AIAA J., 28(12), 2034 (1990). The simulations were conducted using the spectral-element code Nek5000.

Hypersonic Conical Shock-Wave/Turbulent-Boundary-Layer Interaction at High Reynolds Number

A parametric study of hypersonic impinging conical-shock-wave/turbulent-boundary-layer interaction (CSBLI) is carried out at hypersonic high-Reynolds-number conditions (Mach number 6.0, , based on the freestream momentum boundary-layer thickness and wall viscosity) by means of numerical simulation of the Reynolds-averaged Navier–Stokes (RANS) equations, with the eventual goal of establishing wall temperature effects. Comparison with available experimental data shows that RANS is capable of predicting the main features of hypersonic oblique shock-wave/turbulent-boundary-layer interaction, namely, typical size and distribution of the wall properties. A large number of flow cases, especially at high Reynolds number, were computed to examine the scaling of the heat transfer over a wide range of wall temperatures. As expected, the interaction zone of hypersonic CSBLI is reduced as the wall is cooled. A simple power of heat transfer originally introduced by Back and Cuffel (“Changes in Heat Transfer from Turbulent Boundary Layers Interacting with Shock Waves and Expansion Waves,” AIAA Journal, Vol. 8, No. 10, 1970, pp. 1871–1873) for planar shock-induced interactions is here considered to account for hypersonic CSBLIs, which is found to successfully collapse the data to the distributions obtained for supersonic/hypersonic, cold/hot interactions. The value range of the power exponent is within 0.75–0.95.

Defining the Design Space for Double–Double Laminates by Considering Homogenization Criterion

The double–double (DD) laminate families that contain two continuous angles, which were proposed by Tsai (“Double–Double: New Family of Composite Laminates,” AIAA Journal, Vol. 59, No. 11, 2021, pp. 4293–4305), opened up a whole new era for composite layups, which are easy to manufacture and design. In the present study, the design space referred to as feasible regions is derived explicitly based on novel formulations for the lamination parameters of DD laminates. This enables the boundaries of the design space to be obtained analytically, providing mathematical support for DD families. The obtained result shows that their design space is larger than that of conventional quadaxial laminates in terms of industrial practices. A homogenization criterion is implemented into the design space, based on which a tailored DD laminate is proposed, expanding design possibilities and enabling homogenization to be achieved using only 16 plies/4 repeats. The work proposed offers significant benefits through practical solutions to making design, manufacturing, and testing simpler and more competitive.

The effect of nozzle contour on the vibroacoustic loads that form from high area ratio rocket nozzles during sea level launch

An accurate assessment of the vibro-acoustic loads that form during startup of large area ratio rocket nozzles is important for sea-level launch vehicle design and certification. These loads are driven principally by various flow and shock patterns that form inside the nozzle, which are unique to the nozzle contour. This presentation will review a number of laboratory-scale measurements of different nozzle contours and configurations reported by Baars and Tinney, Exp. Fluids, (2013), Donald et al. AIAA Journal (2014), Canchero etal. AIAA Journal (2016), and Rojo et al. AIAA Journal (2016) as it relates to launch platforms of current interest. In particular are the various sources of noise pertaining to transonic resonance, broadband shock associated noise, and the end-effects-regime (EER). The latter of these is unique to the thrust-optimized parabolic contour nozzle as is used on the current Space Launch System vehicle. This EER event occurs when the annular flow structure is in a partial restricted-shock separated (RSS) flow state and is categorized by an onset of relatively low frequency energy driven by intermittent buffeting between RSS flow and partial free shock separated flow at the nozzle lip.

Drag Decomposition Using Partial-Pressure Fields: Ventus-3 and NASA Common Research Model

The development of partial-pressure field (PPF) theory (Schmitz, S “Drag Decomposition Using Partial-Pressure Fields in the Compressible Navier–Stokes Equations,” AIAA Journal, Vol. 57, No. 5, 2019, pp. 2030–2038) and the application of the concept to conditions relevant to commercial transport aircraft, operating in both the subsonic and transonic regimes (Hart, P. L., and Schmitz, S., “Drag Decomposition using Partial-Pressure Fields: ONERA M6 Wing,” AIAA Journal, Vol. 60, No. 5, 2022 pp. 2941–2951), have shown promise as complements to classical far-field methods. In the present work, a further comprehensive analysis of drag decomposition methods is applied to wing and aircraft configurations. First, the Ventus-3 sailplane wing case demonstrates the ability of PPFs to successfully deconstruct drag on industry-relevant cases including transition modeling and winglets. A quantitative comparison to classical aerodynamic theory and far-field methods supports PPF theory and provides additional insight to viscous–inviscid interaction by means of a viscous interactional aerodynamics term in PPF theory. Second, the NASA Common Research Model is simulated in transonic flow to showcase drag decomposition of full aircraft configurations where PPF applications are used to provide further insight into wave drag sources acting on the aircraft in comparison to classical far-field decomposition methods.

Nonintrusive Experimental Aeroelastic Analysis of a Highly Flexible Wing

The aeroelastic response of the Delft–Pazy wing to steady and periodic unsteady inflow conditions is analyzed experimentally. The Delft–Pazy wing is a highly flexible wing model based on the benchmark Pazy wing (Avin, O., Raveh, D. E., Drachinsky, A., Ben-Shmuel, Y., and Tur, M., “Experimental Aeroelastic Benchmark of a Very Flexible Wing,” AIAA Journal, Vol. 60, No. 3, 2022, pp. 1745–1768) and exhibits wingtip displacements of more than 24% of the span in the present study. The nonintrusive measurements are performed with an integrated optical approach that provides combined measurements of the structural response of the wing and the unsteady flowfield around it. The aeroelastic loads acting on the wing are derived using physical models and validated against force balance measurements, showing a good agreement for all considered inflow conditions. The analysis of the aeroelastic response of the wing to the unsteady inflow produced by a gust generator shows that both structural and aerodynamic responses depend strongly on the frequency of the gust. The results of this study provide a characterization of the aeroelastic behavior of the Delft–Pazy wing and can serve as a reference for the development of novel and improved nonlinear aeroelastic simulation models.

Reduced-Order Comparison of Simulated and Measured Coalescing Mach Waves near Supersonic Jets

Prior measurements of the sound field produced by a laboratory-scale, Mach 3 jet flow (Baars and Tinney, Journal of Sound and Vibration, Vol. 333, No. 12, 2014, pp. 2539–2553; Fiévet et al., AIAA Journal, Vol. 54, No. 1, 2016, pp. 254–265) suggest that acoustic waveforms steepen early on in their development. This explained the discrepancy between theoretical predictions, based on effective Gol’dberg numbers, that shocks should not form, and observations of steepened Mach waves close to laboratory-scale jets. The present work continues studying this phenomenon by exploring coalescence processes that occur when neighboring waveforms intersect, forming larger-amplitude waveforms with increased cumulative nonlinear distortion. A numerical model based on the Khokhlov–Zabolotskaya–Kuznetsov (KZK) equation is developed to show that coalescence-induced steepening is sensitive to the intersection angle between adjacent waveforms, waveform duration, and cylindrical spreading effects. High frame-rate schlieren images of sound waves propagating from the post-potential core region of a laboratory-scale Mach 3 jet are then captured along an angle following the ridge of most intense noise to study the development and evolution of coalescence. A shock detection algorithm isolates shock-like events, which are tracked using a translating coordinate system and decomposed using proper orthogonal decomposition. Reduced-order reconstructions of both schlieren images and the KZK model identify common patterns that characterize the shock coalescence process.

Experimental Characterization of an Axisymmetric Transonic Separated Flow for Computational Fluid Dynamics Validation

An experimental characterization of transonic, turbulent, separated flow generated by an axisymmetric model is presented, with intent for use as a computational validation case. The model is a scaled version of a geometry inspired by Bachalo and Johnson (“Transonic, Turbulent Boundary-Layer Separation Generated on an Axisymmetric Flow Model,” AIAA Journal, Vol. 24, No. 3, 1986, pp. 437–442), consisting of a circular bump on a constant-diameter cylinder aligned with the flow. The Mach 0.875 flow is turbulent approaching the bump and becomes locally supersonic at the apex. This leads to a shock-wave/boundary-layer interaction, a separation bubble, and unsteady flow reattachment downstream. Tunnel boundary conditions are characterized, and mean surface pressure, mean skin friction, and both mean and fluctuating velocity fields are measured throughout the interaction region. Uncertainty estimates are provided for all measurements. The degree of flow axisymmetry and evidence of relaminarization are discussed, and a comparison is made to historical data.

Hypersonic Shock Wave/Turbulent Boundary Layer Interaction over a Compression Ramp

A parametric study of ramp-induced planar shock-wave/turbulent-boundary-layer interactions (SBLIs) is carried out at hypersonic conditions (Mach number 6.0) by means of numerical simulation of the Reynolds-averaged Navier–Stokes (RANS) equations, with the eventual goal of establishing wall temperature and Reynolds number effects. Comparison with available experimental data shows that RANS is capable of predicting the main features of hypersonic oblique SBLI, namely, typical size and distribution of the wall-surface pressure, and heat transfer. A large number of flow cases, at low () and high Reynolds number (), were computed to examine the scaling of the heat transfer over a wide range of wall temperatures. As expected, the interaction zone of hypersonic ramp-induced SBLI is reduced as the wall is cooled. A simple power law for heat transfer originally introduced by Back and Cuffel (AIAA Journal, Vol. 8, No. 10, 1970, pp. 1871–1873) is here considered to account for hypersonic ramp-induced SBLI, which is found to successfully collapse the data to the distributions obtained for supersonic, cold/hot interactions.

Empirical Closure Model for Coupling Mode Prediction in Supersonic Rectangular Twin Jets

An empirical closure model of rectangular twin jets’ screech and coupling has been developed to predict their coupling mode at a given flow condition. It builds on Powell’s original screech closure model (Powell, A., “On the Mechanism of Choked Jet Noise,” Proceedings of the Physical Society, Section B, Vol. 66, No. 12, 1953, pp. 1039–1056) of feedback waves produced by the interactions of large-scale structures (LSSs) with each shock cell. Powell’s work is expanded to include nonuniform shock-cell spacing, overexpanded flow regimes, and twin jet configurations. The feedback wave interference pattern strength at the nozzle exit is postulated to determine the realized coupling mode. The required empirical parameters are the shock-cell streamwise locations and the convective velocity of the LSSs. The model was validated using data from an experimental investigation of rectangular twin jets at overexpanded and underexpanded conditions. The coupling mode is correctly predicted for 11 of the 12 nozzle pressure ratios tested. Two datasets from the literature (Jeun et al., “Aeroacoustics of Twin Rectangular Jets Including Screech: Large-Eddy Simulations with Experimental Validation,” AIAA Journal, Vol. 60, No. 11, 2022, pp. 1–21 and Raman, G., and Taghavi, R., “Coupling of Twin Rectangular Supersonic Jets,” Journal of Fluid Mechanics, Vol. 354, Jan. 1998, pp. 123–146) were also examined, and the model made accurate predictions of the coupling mode for all seven reported cases. The response of the twin jets to flow control at various artificially imposed frequencies and excitation modes strongly supports the model predictions, highlighting their usefulness as a guide for the implementation of active flow control in twin jets.

Robustness of Reduced-Order Jet Noise Models

Three statistical jet noise prediction models are compared for a representative set of single-stream jet cases, which include cold and hot jets of the Strategic Investment in Low-Carbon Engine Technology experiment at an acoustic Mach number of 0.875 as well as the cold jets of the NASA small hot jet acoustic rig experiment at acoustic Mach numbers of 0.5 and 0.9. The implemented models are those proposed by Tam and Auriault (Jet Mixing Noise from Fine-Scale Turbulence,” AIAA Journal, Vol. 37, No. 2, 1999, pp. 145–153), Khavaran and Bridges (“An Empirical Temperature Variance Source Model in Heated Jets,” NASA TM 2012-217743, 2012), and the Goldstein’s “A Generalized Acoustic Analogy,” Journal of Fluid Mechanics, Vol. 488, July 2003, pp. 315–333) generalized acoustic analogy (GAA) model. By virtue of reduced-order modeling, which is based on the single-point mean flow and turbulence statistics, all of these implementations use a number of empirical dimensionless source parameters for far-field noise spectra predictions. In comparison with the Tam and Auriault model, the Khavaran and GAA model implementations use several dimensionless parameters, which are available from previous literature and assumed to be more or less universal for a class of single-stream jets. These parameters include the fluctuating enthalpy function and the dimensionless amplitudes of autocovariances of turbulent fluctuating stresses and velocities available from the literature. The comparison of the three models is aimed at not only assessing their accuracy for a range of jet conditions, observer angles, and frequencies but also to examine their robustness outside of a reference jet experiment for which their source models were calibrated. For the input to each model, the mean flow, turbulence kinetic energy, and dissipation rate extracted from Large Eddy Simulations and Reynolds-averaged Navier–Stokes solutions are considered.

Modular Method for Estimation of Velocity and Temperature Profiles in High-Speed Boundary Layers

The fluid property variation caused by viscous heating affects the mean velocity and Reynolds stresses in compressible turbulent boundary layers, and as a result also affects the resulting skin friction and wall heat transfer. We present a method based on the work of Huang et al. (AIAA Journal, Vol. 31, No. 9, 1993, pp. 1600–1604) that estimates the mean velocity and temperature profiles, and therefore also the friction and heat transfer coefficients, for a given Mach number, Reynolds number, and wall thermal condition. The method is modular in the sense that it works with multiple variable-property scaling formulas (or “transformations”), velocity–temperature relationships, viscosity–temperature relationships, and equations-of-state. The method is validated against multiple Direct Numerical Simulation (DNS) cases and was found to predict the skin friction and wall heat flux to within 8 and 11% compared to the current state-of-the-art (the Van Driest II method), which is within 16 and 13% for the same data.

Flutter Tests of the Pazy Wing

The paper presents the flutter and limit-cycle oscillation (LCO) tests of the Pazy wing, a very flexible wing model designed to study aeroelastic phenomena associated with nonlinear geometry and provide data for nonlinear aeroelastic model validation. The Pazy wing model, analyses, and data from earlier aeroelastic tests are provided by Avin et al. (“Experimental Aeroelastic Benchmark of a Very Flexible Wing,” AIAA Journal, 2022, pp. 1–24). In the current study, three velocity sweeps were performed at different angles of attack (AoA), during which the wing entered and exited limit-cycle oscillations (LCO). A motion-recovery camera system and fiber-optics Bragg grating sensors were used to track the deformations and strains over the wingspan. The wing encountered LCO when the wingtip deformation was approximately 25% of the span. The onset velocity, frequency, and mode shape varied, depending on the test’s AoA. The paper first focuses on the onset of the oscillations, where the onset velocity and oscillation mode-shapes are analyzed. The paper then continues to investigate the LCO characteristics and the nonlinear behavior of the wing. The wing models and experimental data are publicly available through the Third Aeroelastic Prediction Workshop.**

Pitch/Plunge Equivalence for Dynamic Stall of Unswept Finite Wings

Pitch/plunge equivalence for deep dynamic stall of a straight finite wing is considered by employing large-eddy simulations. The flowfields are computed with a well-established time-implicit high-fidelity large-eddy simulation approach. The wing has a moderate aspect ratio of and a NACA 0012 section incorporating a rounded tip. The flow parameters are a freestream Mach number of and chord Reynolds numbers of . Pure pitching and equivalent plunging maneuvers are considered with a reduced frequency of and minimum and maximum effective angles of attack of 4 and 22 deg, respectively. The pitch and plunge dynamic stall cases are compared in detail in terms of the unsteady three-dimensional flow structure, surface pressure, and loading. Differences in the lift and moment coefficient are shown to be correctable, employing a steady thin-airfoil theory for a cambered airfoil that accounts for the rotation-induced apparent camber present in the pitching case. This pitch/plunge equivalence is extendable to the dynamic stall of a swept wing, as demonstrated in another paper by the authors [“Pitch–Plunge Equivalence for Dynamic Stall of Swept Finite Wings,” AIAA Journal (submitted for publication)].

Pitch/Plunge Equivalence for Dynamic Stall of Swept Finite Wings

A high-fidelity numerical study is undertaken to expand the notion of pitch/plunge equivalence to finite, swept wings undergoing deep-dynamic stall and builds upon the analysis of straight wings in another paper by the authors [“Pitch/Plunge Equivalence for Dynamic Stall of Unswept Finite Wings,” AIAA Journal (submitted for publication)]. The wing has an aspect ratio of with a NACA 0012 section, a rounded tip, and a sweep angle of ; it operates at a Reynolds number of and a Mach number of . Equivalent pure-pitch and pure-plunge motions at two different reduced frequencies are considered to assess pitch-induced apparent camber on the wing and its effects on the three-dimensional flow structure, surface topology, and aerodynamic loading. Regardless of the motion type or rate, the wing undergoes dynamic tip stall that is initiated through the outboard separation of a dynamic stall vortex as it lifts off the surface and forms into an arch vortex before ejecting into the wake. The pitching cases exhibit an advanced development and evolution in time of these features that are observed as variable amplifications of the aerodynamic loading, which is due predominantly to an apparent camber effect. Corrections to the total and sectional lift and moment coefficients are explored by extending steady thin-airfoil theory to pitching swept wings. Remarkably, the corrected plunge-equivalent histories provide a full collapse of the force and moment responses between the pitching and plunging cases, despite the delayed behavior of the flow structure. This equivalence and the similar process of dynamic stall should facilitate the comparison of experiments and/or simulations employing different motion types, even for finite-aspect-ratio and swept wings.

Wake of a Coaxial Corotating Rotor in Hover

A recent paper by Tinney and Valdez (AIAA Journal, Vol. 58, No. 4, 2020, pp. 1657–1667) reports the trade space between near-field acoustics and thrust produced by a laboratory-scale, coaxial, corotating rotor in hover by considering the effect of various index angles and stacking distances between the upper and lower rotor for a range of rotor speeds. The current effort is a follow-up campaign employing high-speed digital schlieren and single camera particle image velocimetry to measure the flowfield produced by the same stacked rotor. Particle image velocimetry measurements quantify the evolutionary behavior of the upper and lower blade tip vortices, and their interactions for various index angles and wake ages for a given rotor speed. Vortex tracking methods quantify the anisotropic wandering motions of the blade tip vortices and reveal patterns that are unique to both the upper and lower rotors. The findings reinforce the notion that the optimal index angle for the thrust produced by a coaxial, corotating rotor is one where the blade tip vortices produced by the upper rotor graze the low-pressure side of the lower rotor. If the miss distance is too small, collisions between the upper rotor vortex and the lower rotor ensue so that any gains in thrust are compounded by blade-vortex-interaction noise penalties.

Calibration of the Transitional Turbulence Model

In the present paper, the turbulent transition model from Langtry and Menter (“Correlation-Based Transition Modeling for Unstructured Parallelized Computational Fluid Dynamics Codes,” AIAA Journal, Vol. 47, No. 12, 2009, pp. 2894–2906) is implemented in the structured finite volume Reynolds-averaged Navier–Stokes code FANSC and coupled with the shear-stress transport turbulence model. The model is modified and recalibrated to better match experimental results. The local turbulent intensity is replaced with a more consistent turbulent fluctuation representation to avoid nonphysical solutions at the stagnation point that would contaminate the downstream boundary-layer predictions. A calibration is devised by customizing several constants involved in the equation for the intermittency and the destruction term for the turbulent kinetic energy . An optimization algorithm is thus employed to facilitate the calibration on well-known two-dimensional test cases, such as the flat plates of the T3 series and of Schubauer and Klebanoff (“Contributions on the Mechanics of Boundary-Layer Transition,” NACA TN 3489, 1955), as well as the NACA 0012 airfoil. The calibrated model is then validated on the NLF-0416 and S809 airfoil cases that have not been involved in the optimization process.