Wide Range Of Mach Numbers Research Articles

A transonic, viscous nonlinear frequency domain Vortex Lattice Method for aeroelastic analyses

The Unsteady Vortex Lattice Method has interesting capabilities for nonlinear subsonic aeroelastic analyses. This paper aims to develop an aeroelastic framework based on the UVLM that is suitable for compressible, transonic and viscous unsteady flows. To achieve these modeling capabilities, the UVLM is first extended to the nonlinear frequency domain using the harmonic balance approach. Secondly, sectional corrections are applied from a steady 2D/2.5D database of aerodynamic forces computed with high-fidelity Euler/RANS flow solvers. The angle of attack coupling scheme on the lift force is complemented by a least-square procedure also coupling the pitching moment. Finally, an approximation of the retarded time is implemented in the influence kernel of the harmonic balance formulation of the UVLM to account for unsteady compressibility. The aerodynamic model is verified against time domain high-fidelity solutions with pitching airfoils in 2D and in 3D with a pitching constant chord NACA 0012 swept wing. Thereafter, the HB/LCO approach is applied with the frequency domain UVLM to 2D and 3D aeroelastic problems and compared against high-fidelity solutions. The proposed framework is capable of capturing strong nonlinear Limit Cycle Oscillation behaviors at high Mach numbers, as well as the flutter boundary for a wide range of Mach numbers while maintaining a low computational cost. As a result, the proposed framework could be of interest for preliminary design and Multi-Disciplinary Design (MDO) optimizations in the context of subsonic and transonic aircraft design.

Very-high-order TENO schemes with adaptive accuracy order and adaptive dissipation control

In this paper, a new family of very-high-order TENO schemes with adaptive accuracy order and adaptive dissipation control (TENO-AA) is proposed. The new framework allows for constructing arbitrarily high-order TENO schemes in a unified paradigm and the yielded nonlinear schemes gradually reduce to low-order reconstructions by judging the smoothness with the ENO-like stencil selection strategy. In order to control the nonlinear numerical dissipation adaptively, the flow scales are first measured by examining the first-order undivided difference and the cut-off constant CT in the TENO weighting strategy is adapted based on the corresponding measurement. With one set of optimal parameters, the newly proposed TENO schemes are designed to deliver excellent performance for predicting highly compressible flows with a wide range of Mach numbers. While the new very-high-order TENO schemes feature good robustness for conventional gas dynamics, the ENO-property is well preserved with the assistant of a positivity-preserving flux limiter for extreme simulations. Without loss of generality, the typical eight- and ten-point TENO-AA schemes are constructed. A set of benchmark simulations are computed to demonstrate the performance of the proposed TENO schemes.

A low-dissipation HLLD approximate Riemann solver for a very wide range of Mach numbers

We propose a new Harten-Lax-van Leer discontinuities (HLLD) approximate Riemann solver to improve the stability of shocks and the accuracy of low-speed flows in multidimensional magnetohydrodynamic (MHD) simulations. Stringent benchmark tests verify that the new solver is more robust against a numerical shock instability and is more accurate for low-speed, nearly incompressible flows than the original solver, whereas additional computational costs are quite low. The novel ability of the new solver enables us to tackle MHD systems, including both high and low Mach number flows.

A Surrogate-Based Method for Compressible Boundary-Layer Transition Prediction

To improve robustness and efficiency of automatic transition prediction in aerodynamic design, a reduced model of linear stability analysis is usually adopted, such as -envelope or -database method. Nevertheless, building such a model is challenging when it comes to compressible flows, as the transition mechanism is more complex and multiple flow parameters should be taken into consideration. To address this problem, this paper proposes an efficient surrogate-based method for compressible boundary layers that uses pretrained surrogate models to substitute linear stability analysis, concerning stability analysis of both Tollmien–Schlichting and Mack modes, as well as transition prediction of flow over arbitrary-shaped airfoils. The proposed method is demonstrated by stability analysis of compressible flat-plate boundary layers at a wide range of Mach numbers of . It is also validated by transition prediction of flow over a low-speed natural-laminar-flow (NLF) airfoil NLF-0416 and a transonic NLF airfoil NPU-LSC-72613. Besides, a sample partitioning method is presented to accelerate surrogate-model training with large samples. Results show that the predicted growth rates of perturbations, factors, and corresponding transition locations by our method of using surrogate-based stability analysis agree well with those by a standard method of solving full linear stability equations.

Coagulation of inertial particles in supersonic turbulence

Coagulation driven by supersonic turbulence is primarily an astrophysical problem because coagulation processes on Earth are normally associated with incompressible fluid flows at low Mach numbers, while dust aggregation in the interstellar medium for instance is an example of the opposite regime. We study coagulation of inertial particles in compressible turbulence using high-resolution direct and shock-capturing numerical simulations with a wide range of Mach numbers from nearly incompressible to moderately supersonic. The particle dynamics is simulated by representative particles and the effects on the size distribution and coagulation rate due to increasing Mach number is explored. We show that the time evolution of particle size distribution mainly depends on the compressibility (Mach number). We find that the average coagulation kernel ⟨Cij⟩ scales linearly with the average Mach number ℳrms multiplied by the combined size of the colliding particles, that is, 〈Cij〉∼〈(ai+aj)3〉 ℳrmsτη−1, which is qualitatively consistent with expectations from analytical estimates. A quantitative correction 〈Cij〉∼〈(ai+aj)3〉(vp,rms/cs)τη−1 is proposed and can serve as a benchmark for future studies. We argue that the coagulation rate ⟨Rc⟩ is also enhanced by compressibility-induced compaction of particles.

Rapid airfoil design optimization via neural networks-based parameterization and surrogate modeling

Aerodynamic optimization based on computational fluid dynamics (CFD) is a powerful design approach because it significantly reduces the design time compared with the human manual design. However, CFD-based optimization can still take hours to converge because it requires repeatedly running computationally expensive flow simulations. To further shorten the design optimization time, we propose a fast, interactive design framework that allows us to complete an airfoil aerodynamic optimization within a few seconds. This framework is made efficient through a B-spline-based generative adversarial network model for shape parameterization, which filters out unrealistic airfoils for a reduced design space that contains all relevant airfoil shapes. Moreover, we use a combination of multilayer perceptron, recurrent neural networks, and mixture of experts for surrogate modeling to enable both scalar (drag and lift) and vector (pressure distribution) response predictions for a wide range of Mach numbers (0.3 to 0.7) and Reynolds numbers (104 to 1010). To verify our proposed framework, we compare the optimization results with the ones computed by direct CFD-based optimization for subsonic and transonic conditions. The results show that the optimal designs and the aerodynamic quantities (lift, drag, and pressure distribution) obtained by our proposed framework agree well with the ones computed by direct CFD-based optimizations and evaluations. The proposed framework is being integrated into a web-based interactive aerodynamic design framework that allows users to predict drag, lift, moment, pressure distribution, and optimal airfoil shapes for a wide range of flow conditions within seconds.

An integration method based on a novel combined flow for aerodynamic configuration of strutjet engine

In this paper a novel design method of aerodynamic configuration is proposed to integrate forebody, strut and inlet for strutjet engine, and a model at design point of Mach number 6 is generated to investigate the aerodynamic performance by both simulations and experiments. The basic flow field employed by proposed method is a combined flow named IBB, which is combined by Internal Conical Flow A (ICFA), truncated Busemann flow I (BI) for external section, and truncated Busemann flow II (BII) for internal section. The model configuration is generated by streamline tracing method from basic flow field, in which the forebody section is traced from ICFA and BI flows, and the inlet as well as strut section is traced from BII flow. The simulations in Mach number 4, 5, and 6 demonstrate uniform starting flow fields with relatively high total pressure recovery, which agree well with experiments in wind tunnel. Additionally, in low Mach number cases, this inlet could start at Mach number 3 while it is unstarted at Mach number 2.7; in high Mach number cases, a uniform flow could still exist in Mach number 6.5 while a relatively strong shock wave boundary layer interaction is found in cowl area of Mach number 7 case, indicating the inlet designed by proposed method works in a relatively wide Mach number range.

Higher Order Sliding Mode-Based Guidance for Controlling Statically Unstable Missiles

A surface to air missile (SAM) typically operates at a wide range of Mach numbers, angles of attack and altitudes. This leads to an inevitable significant variation in center of pressure (CP) location. In addition, the propulsion burning changes the Center of Gravity (CG) location. The variation in CP and CG locations results in high variation in aerodynamic static stability in the flight envelope. The high variation of static stability makes operating at statically unstable regions inevitable and needs to be tackled through the missile on-board algorithm. The control design poses a requirement of high autopilot bandwidth to tackle the static instability, which leads to unrealistic actuator bandwidth requirements. Here, we propose a novel technique to handle the high static instability outside the purview of control design without burdening the control requirements. In this paper, the highly unstable operating conditions are avoided through judicious design of mid-course guidance. This paper also presents design of this guidance scheme through higher order sliding mode technique.

STREAmS: A high-fidelity accelerated solver for direct numerical simulation of compressible turbulent flows

We present STREAmS, an in-house high-fidelity solver for direct numerical simulations (DNS) of canonical compressible wall-bounded flows, namely turbulent plane channel, zero-pressure gradient turbulent boundary layer and supersonic oblique shock-wave/boundary layer interaction. The solver incorporates state-of-the-art numerical algorithms, specifically designed to cope with the challenging problems associated with the solution of high-speed turbulent flows and can be used across a wide range of Mach numbers, extending from the low subsonic up to the hypersonic regime. From the computational viewpoint, STREAmS is oriented to modern HPC platforms thanks to MPI parallelization and the ability to run on multi-GPU architectures. This paper discusses the main implementation strategies, with particular reference to the CUDA paradigm, the management of a single code for traditional and multi-GPU architectures, and the optimization process to take advantage of the latest generation of NVIDIA GPUs. Performance measurements show that single-GPU optimization more than halves the computing time as compared to the baseline version. At the same time, the asynchronous patterns implemented in STREAmS for MPI communications guarantee very good parallel performance especially in the weak scaling spirit, with efficiency exceeding 97% on 1024 GPUs. For overall evaluation of STREAmS with respect to other compressible solvers, comparison with a recent GPU-enabled community solver is presented. It turns out that, although STREAmS is much more limited in terms of flow configurations that can be addressed, the advantage in terms of accuracy, computing time and memory occupation is substantial, which makes it an ideal candidate for large-scale simulations of high-Reynolds number, compressible wall-bounded turbulent flows. The solver is released open source under GPLv3 license. Program summaryProgram Title: STREAmSCPC Library link to program files:https://doi.org/10.17632/hdcgjpzr3y.1Developer’s repository link:https://github.com/matteobernardini/STREAmSCode Ocean capsule:https://codeocean.com/capsule/8931507/tree/v2Licensing provisions: GPLv3Programming language: Fortran 90, CUDA Fortran, MPINature of problem: Solving the three-dimensional compressible Navier–Stokes equations for low and high Mach regimes in a Cartesian domain configured for channel, boundary layer or shock-boundary layer interaction flows.Solution method: The convective terms are discretized using a hybrid energy-conservative shock-capturing scheme in locally conservative form. Shock-capturing capabilities rely on the use of Lax–Friedrichs flux vector splitting and weighted essentially non-oscillatory (WENO) reconstruction. The system is advanced in time using a three-stage, third-order RK scheme. Two-dimensional pencil distributed MPI parallelization is implemented alongside different patterns of GPU (CUDA Fortran) accelerated routines.

A quantitative study of the aerodynamic characteristics of a low aspect ratio aircraft when flying at near-critical Mach numbers

This paper presents calculations of aerodynamic characteristics for an aircraft of the conventional aerodynamic design, i.e. with a wing module of the X-type and a tail assembly of the cruciform design. The calculations were performed using a numerical simulation method based on Reynolds-averaged Navier – Stokes equations. Integral aerodynamic coefficients were calculated within a wide range of Mach numbers. The obtained aero-dynamic characteristics were compared with those of aircrafts with a rectangular wing and a trapezoidal wing. The contribution of outer wings in the generation of ascensional power was assessed. The determination of the aerodynamic characteristics of an aircraft in the range of near-critical Mach numbers is necessary both for mathematical modelling of control systems and for establishing possible flight modes.

Controllable initiation characteristics of the oblique detonation wave in a combustor with a confined cone of a novel structure

The reliable and controllable initiation of the oblique detonation wave at an appropriate location is vital to maintain the optimal thrust of the propulsion systems. A novel structure is proposed in the current research to dynamically control the initiation of the oblique detonation wave in a combustor with a confined cone. The novel structure is based on the traditional single cone, of which the cone angle is θ0. The novel cone utilizes a leading cone with a smaller cone angle (θ1) + a larger rear cone angle (θ2), where θ1<θ0<θ2. A parametric study is conducted based on reactive Euler equations to test the performance of the novel structure on detonation initiation. It is found that the oblique detonation can be stabilized with less total pressure loss and less drag force under a wide range of Mach numbers. It is found that an ‘internal reflection’ of the rear shock wave exists in the first and second reflection of the leading shock wave. The ‘internal reflection’, which has not been reported before, enhances the temperature elevation and pressure elevation in the induction region of the oblique detonation wave. A possible strategy of dynamic control based on the novel structure is proposed. The force exerted on the rear cone is different under different Mach numbers, pushing the rear cone to a state of a certain cone angle. The novel structure performance very well for controllable initiation of the oblique detonation wave.

Closed-loop flow control of an ultra-compact serpentine inlet based on nondimensional model

Ultra-compact serpentine inlet faces serve inlet-engine compatibility issues due to flow distortion. To ensure inlet-engine compatibility over a wide range of Mach number, novel active flow control techniques with the ability of being opened or adjusted as needed draw many attentions in recent years. In this paper, a feedback control system was developed based on the method of microjet blowing. The proposed system includes a pressure adjusting valve to adjust the control effort, a dynamic pressure sensor to sense the inlet distortion intensity, a signal processing instrument to calculate the Root-Mean-Squared (RMS) pressure, and a controller to implement feedback control. To achieve high quality closed-loop controls at dynamic conditions, a novel nondimensional feedback method was developed. The advantage of this nondimensional method was validated at both off-design and arbitrarily changing Mach number conditions. With a sectional PI control law, the RMS control error reduced more than 56% at arbitrary changing conditions. Works in this paper also showed that the dynamics of this nondimensional system can be simplified as a stable second-order overdamped system.

Design of Diaphragmless Shock Tube and Research on its Normal Temperature Characteristics

Shock wave is an important physical phenomenon in supersonic airflow. The shock tube is an experimental device that generates shock wave. It generates shock wave by generating a pressure difference between its high-pressure section and low-pressure section. Most researchers in the world use diaphragm shock tubes to generate shock wave, such as single diaphragm structure and double diaphragms structure type shock tubes. Although the type of shock tube can generate an incident shock wave with a wide range of Mach numbers, in the process of producing shock wave and breaking film, there will be broken film fragments interfering with the flow field. And every time the researcher needs to replace the diaphragm after the experiment, the efficiency of the experiment is low. In this paper, a diaphragmless shock tube is designed. The high-speed reciprocating motion of two pistons in the high-pressure section is used to replace the process of membrane rupture. And the normal temperature characteristics of the shock tube are researched.

Average Friction Factors of Choked Gas Flow in Microtubes

A micro-tube passage is a basic and important element in the design of micro heat exchangers and for this reason during the last decade a series of investigations have been made with the aim to clarify the main scaling effects playing an important role in microtubes. In this paper, a combined analysis of numerically and experimentally obtained average friction factors in microtubes under the situation of under-expanded (choked) gas flow is presented. The working fluid (nitrogen) passes through the microtube and discharges into the atmosphere under an increasing inlet pressure. Experiments and numerical computations are performed for microtubes with 249 and 528.9 mm in diameter, by varying the aspect ratio (i.e. length/diameter) from 100 to 200. The numerical methodology to solve the governing equations is based on the Arbitrary-Lagrangian-Eulerian (ALE) method. In order to capture the under-expansion characteristics of the flow during choking, the computational domain is extended in the downstream region beyond the microtube outlet. Both experimental and numerical results were obtained for a wide range of Mach number and Reynolds number. In the previous study, it was demonstrated how the outlet Mach number can be expressed as a function of the tube diameter under choked conditions. In this paper, a data reduction procedure for the estimation of the average friction factor between the inlet and the outlet of the microtube is proposed for choked flows in which the outlet gas temperature and pressure are obtained by using the outlet Mach number calculated numerically as a function of the microtube diameter. It is demonstrated how this data reduction method allows an accurate calculation of the average friction factors in microtubes by using a limited number of parameters which are easy to measure. The results obtained in this way are in good agreement with the numerical predictions as well as with the most common empirical correlations.

Semi-implicit method to solve compressible multiphase fluid flows without acoustic time step restrictions

A new 3D multiphase numerical capability is presented here for simulating multiphase flow regimes at all Mach numbers (M). The new method is a semi-implicit extension of the finite volume discrete equation method (DEM) of Chinnaya et al. (2004), which originally used explicit time-stepping. The capability is also developed to work with another extension of the DEM to moving grids for arbitrary Lagrangian-Eularian (ALE) methods detailed in Dunn (2011). Rather than solving all phase equations simultaneously, the DEM reduces the equations to a system of single-phase Riemann solves, where each phase has its own velocity and thermodynamic state. Exchanges between the phases are modeled through source terms accounting for the phase interactions. Since the original multiphase scheme uses an explicit time-advancement scheme, it has time step restrictions dictated by the speed of sound, which limits the model's ability to simulate weakly compressible flows which typically need to be integrated for longer time periods. Here, we extend the current multiphase formulation by implementing a pressure-correcting step to enable implicit calculations and remove acoustic time step limitations. The new semi-implicit algorithm allows use of relatively large time steps compared to an explicit method. Validation and benefits of the new implicit time-step method are illustrated using several examples including weakly compressible flows and strong shock waves. The scheme presented here is general and may be used for a variety of applications which require capabilities for handling multiphase flow at a wide range of Mach numbers. However, the goal of this research is to simulate all stages of high energy explosions, including the shock formation (high Mach numbers) and evolution of the buoyant cloud (low Mach numbers).

A sharp-interface immersed boundary method for moving objects in compressible viscous flows

Sharp-interface Immersed Boundary Methods for moving solids in compressible viscous flows often exhibit spurious noise propagating from the moving boundary. In compressible flows, filtering or upwinding discretization techniques are often used to overcome the problem. In the present work, we show that using a conservative energy-preserving finite difference method for convection in combination with a Ghost-Point-Forcing-Method (GPFM) is able to keep controlled the pressure-velocity spurious noise. In order to deal with compressible flow in a wide range of Mach numbers the shock-dynamics appearing in high-speed flows conditions has been addressed hybridising the latter scheme with a fifth-order weighted-essentially-non-oscillatory (WENO) scheme. The latter, in the path of keeping the numerical dissipation minimal, has been confined around the shock locations using a proper detector. In this respect, the method appears a suitable alternative for direct and large-eddy simulation of moving objects in compressible flows. The entire methodology was found to be robust ranging from weakly compressible to highly supersonic flows including shocks. In order to prove the cleanliness and the robustness of the method, several well-documented benchmarks and test cases, in a wide range of both Reynolds and Mach numbers, are presented.

Konvergencija rezultata transoničnih aerotunelskih ispitivanja standardnog modela AGARD-B

AGARD-B is a widely-used configuration of a standard wind tunnel model. Beside its originally intended application for correlation of data from supersonic wind tunnel facilities, it was tested in a wide range of Mach numbers and, more recently, used for assessment of wall interference effects, validation of computational fluid dynamics codes and validation of new model production technologies. The researchers and wind tunnel test engineers would, naturally, like to know the "true" aerodynamic characteristics of this model, for comparison with their own work. Obviously, such data do not exist, but an estimate can be made of the dispersion of test results from various sources and of the probable "mean" values of the aerodynamic coefficients. To this end, comparable transonic test results for the AGARD-B model at Mach numbers 0.77, Mach 1.0 and Mach 1.17 from six wind tunnels were analyzed and average values and dispersions of the aerodynamic coefficients were computed.

Hypersonic plasma flows created by a pulse capillary discharge

The possibility of producing by a pulsed capillary discharge the quasi-stationary high-speed plasma flows in a wide range of Mach numbers (M=3-20), electron number densities (ne∼1011-1015 cm-3) and temperatures (Te∼3-10 eV), rotational and vibrational temperatures (Tr=3800-6000 K and Tv=9000-15000 K), which can be used for setting up plasma aerodynamic experiments on small laboratory facilities, is shown. The most suitable for this purpose is a mode close to supersonic outflow into a vacuum, the transition to which occurs when the total pressure ratio exceeds N = p0/p∞ >5·103. This mode permits to obtain wide-aperture flows (D >10 cm) with spatially extended weakly inhomogeneous segments sufficient to accommodate tested bodies of a centimeter scale.

Aeroacoustic characteristics of subsonic flow from axisymmetric pipe-cavities

A detailed experimental study on the acoustic characteristics of an axisymmetric pipe-cavity jet is conducted for different ratios of cavity depth to length, over a range of subsonic Mach numbers. Power spectral analysis and Scalogram are implemented to unravel the physics of the pipe-cavity resonance. Proper Orthogonal Decomposition analysis is used to understand the flow dynamics of the pipe-cavity jet. The results show that the pipe-cavity resonates close to the first tangential mode for a wide range of Mach numbers. However, an increase in the depth and Mach number leads to the onset of combined resonance modes. Experimentally obtained resonance frequencies are validated against numerical simulations and theoretical predictions. Scalogram results show the existence of nonlinear interactions and mode shifting of resonance modes. A parametric space plot demarcates the domains of operation of various oscillation modes, over the relevant parametric range. The results reveal that the upstream cavity significantly affects the downstream jet flow dynamics, and its far-field noise radiation. Deeper cavities have a stronger influence on the far-field acoustics compared to shallow cavities.

Использование гибридного метода крупных частиц для расчета течений многокомпонентных газовых смесей

Статья посвящена обобщению гибридного метода крупных частиц для численного моделирования течений многокомпонентных газовых смесей при наличии границ раздела газов с различными термодинамическими свойствами. Метод относится к алгоритмам сквозного расчета разрывов. Разностная схема является консервативной, однородной и имеет второй порядок аппроксимации по пространству и времени на гладких решениях. Результаты проверки на тестовых задачах в широком диапазоне чисел Маха и отношений газодинамических параметров подтвердили работоспособность метода. Выполнен анализ численных ошибок в окрестности контактных разрывов на сетках различного разрешения, свидетельствующий о сходимости результатов расчета к автомодельным решениям. This paper is devoted to a generalization of a hybrid large-particle method for the numerical simulation of multicomponent gas mixture flows in the presence of gas interfaces with various thermodynamic properties. The method belongs to the class of shock-capturing and interface-capturing algorithms. The employed difference scheme is conservative and uniform and possesses the second order approximation in space and time on smooth solutions. The obtained numerical results show the efficiency of the method in a wide range of Mach numbers and ratios of gas dynamic parameters. The error analysis performed near the contact discontinuities on grids of various resolutions confirms the convergence of numerical results to the self-similar solutions.