Supersonic Research Articles

Fatigue-induced fracture failure of acrylic-polycarbonate laminated aircraft canopy

The fatigue-induced fracture failure of the aircraft canopy occurred in the poly(methyl methacrylate) (PMMA) layer laminated on polycarbonate (PC) during flight. For more than 24 years, the aircraft had been operated at high altitudes and supersonic flight. To identify the root cause and the mechanism for the formation of the fracture, the fracture surfaces were investigated. The fracture morphologies were characterized using optical microscope (OM) and scanning electron microscope (SEM).In macroscopic observations, the main crack showed a total length of approximately 1.6 m from the front to the right of the crack stop groove when viewed from the front of the canopy. The main crack ran about 0.9 m including partly curved line from the front part to the upper middle one and then reached about 0.7 m in a straight line perpendicular to the right of the crack stop groove. There were two crack ends in the main crack: one was at the lower part of the front, the other was at the right end of the crack stop groove. Numerous macro-cracks visible to the naked eye were distributed only on the front surface of the canopy.In microscopic examination, the voids on the front surface of the outer PMMA layer were formed by the friction heat with air during the supersonic flight. The voids served as the origins, the actual starting point of the crack. The voids slowly grew to macro-cracks vertically or horizontally by the thermal stress during flying at high altitudes. Cracks proceeded in the direction of 90 degrees while being bisected in V-shapes downward from the surface of the PMMA layer with the action of thermal tension. The crack growth represents the typical characteristics of the fatigue crack: multi-origins, ratchet marks, and beach marks. The main crack grew further, forming a slight curved line by connecting adjacent macro-cracks arranged in an almost vertical direction. When crack growth reached a critical point, the catastrophic fracture progressed rapidly from the primary origin of the fatigue crack to both ends due to the action of lateral force. Fast crack zones on both sides showed the same dimple and river patterns.This study explains that the combined and synergistic interaction of the fatigue crack and environmental stresses iteratively occurred on the front surface in the outer PMMA layer of the aircraft canopy due to the continual exposure to high altitudes and supersonic flights, consequently resulting in the fatigue-induced fracture failure.

Enhancing improved advection upstream splitting method on triangular grids: A hybrid approach for improved stability and accuracy in compressible flow simulations

This paper introduces NAUSM+M+AUFS (New Improved Advection Upstream Splitting Method Plus Artificially Upstream Flux Vector Splitting), a novel hybrid computational scheme for simulating compressible flows on triangular grids. The AUSM+M (Improved Advection Upstream Splitting Method) method is enhanced through two key modifications to boost numerical stability and robustness in high Mach number and hypersonic flows. The first modification redefines the interfacial numerical sound velocity, reducing shock anomalies and improving shock-capturing by integrating velocity and characteristic sound speed parameters. The second modification addresses the insufficiency of the pressure flux dissipation term at supersonic speeds by introducing a formulation that increases dissipation proportionally to the Mach number, thereby enhancing performance in high-speed flows. These enhancements constitute the NAUSM+M method. The NAUSM+M+AUFS scheme combines the strengths of NAUSM+M and AUFS (Artificially Upstream Flux Vector Splitting) methods, particularly in overcoming the limitations of NAUSM+M in handling shock instabilities and the carbuncle phenomenon on structure triangular grids. A dynamic switching function adjusts the weighting between NAUSM+M and AUFS, optimizing accuracy and stability based on local flow conditions. Numerical tests demonstrate that NAUSM+M+AUFS significantly outperforms AUSM+M, NAUSM+M, and AUFS, effectively eliminating the carbuncle phenomenon and providing smooth shock wave contours. In steady flow analysis, the new hybrid method achieves convergence speeds comparable to AUFS and shows 15% to 45% superior convergence accelerating than AUSM+M, depending on the convergence rate. In addition, in steady flow analysis, the accuracy of NAUSM+M+AUFS is 46% better than that of AUFS. This approach represents a significant advancement, offering a robust, accurate, and efficient solution for high-speed aerodynamic simulations, with broad applicability across various compressible flow challenges.

Ultrasonic and conventional fatigue behavior, strain rate sensitivity, and structural design methods for wrought and cold spray Al-6061

Cold spray is an additive manufacturing process that accelerates powder particles to supersonic speeds to create repairs and bulk depositions with fine-grained microstructures, high density, and good mechanical properties. Fatigue property measurement for these novel materials is critical for their use in safety-critical components, which can be accelerated with the use of ultrasonic fatigue testing. In this work, ultrasonic (20 kHz) and conventional (20 Hz) fatigue studies were conducted on as-sprayed bulk Al-6061 and conventional wrought Al-6061-T6. Complementary fatigue studies of surface preparation (surface finish and residual stress) and fatigue specimen geometry (round versus flat), as well as hole-drilling residual stress measurements, were undertaken to minimize the influence of these confounding variables. Cold spray Al-6061 exhibits fatigue frequency sensitivity, whereas the wrought material does not. Tensile testing at varied strain rates indicates that a portion of the fatigue frequency effect can be attributed to strain rate sensitivity. Fractographic studies show that crack initiation occurs from unbonded powder particles at the surface at high stress amplitude, and transitions sub-surface at lower stress amplitude. The results of these studies were used to create frequency-corrective models of S-N data and Kitagawa-Takahashi diagrams that can be used to design for fatigue crack initiation and growth resistance.

Hypersonic Flow past LD-Haack Series, Parabolic Series, Power Series Nose Cone: A Comparative Study

This research paper presents a detailed aerodynamic analysis of three rocket nose cone designs—LD-Haack, Parabolic, and Power Series—under hypersonic conditions ranging from Mach 5 to Mach 12. Using advanced Computational Fluid Dynamics (CFD) simulations in ANSYS Fluent, the study focuses on key aerodynamic parameters, including drag force, drag coefficient, velocity, pressure, and temperature distributions. The results indicate that the Parabolic Nose Cone offers the most favorable aerodynamic performance, with the lowest drag force and drag coefficient across all Mach numbers. This superior performance can be attributed to its ability to streamline airflow and minimize boundary layer separation, effectively reducing pressure drag and limiting the onset of turbulent wake regions. In contrast, the LD-Haack Series Nose Cone, though designed to minimize wave drag at supersonic speeds, shows moderate drag at hypersonic velocities due to its less optimal control of boundary layer behavior. The Power Series Nose Cone, despite its highly tapered shape, experiences the highest drag. This is largely caused by increased flow separation and turbulent wake generation, which exacerbate pressure drag, particularly at higher Mach numbers. These findings highlight the critical influence of nose cone geometry on aerodynamic efficiency and emphasize the importance of optimizing designs for hypersonic applications. This study provides valuable insights for aerospace engineers seeking to reduce drag and improve performance in high-speed flight scenarios.

A comparative analysis of the aerodynamic performance of supersonic missiles with conical and ogive nose shapes

This paper presents a numerical study conducted to analyze the aerodynamic performance of supersonic missiles consisting of a cylindrical body and four flat-plate rear fins arranged uniformly, equipped with conical and ogive heads. Computational Fluid Dynamics (CFD) simulations were performed using the ANSYS Fluent 17.1 solver, along with the Gambit grid generation software. The objective was to compare the aerodynamic characteristics of these two head designs in terms of drag, lift, and stability at supersonic speeds. Various flow parameters, including Mach number and angle of attack, were investigated to comprehensively assess the performance of the missile configurations. The results indicate clear differences in the aerodynamic behavior of conical and ogive heads. Specifically, there was a 2–11 percent increase in the lift coefficient of the conical heads compared to the ogive heads, and an increase in the drag coefficient of both conical and ogive heads.

Optimization and Analysis of Flow Field Simulations Across Different Mach Regimes Using Discrete Numerical Methods

This article explores the development of flow field models in steady-state environments utilizing Euler equations and potential flow equations, with verification processes conducted using Python. The models demonstrate stability and high accuracy in low-velocity scenarios, capturing essential dynamics effectively. However, as the conditions transition to supersonic speeds, the models begin to exhibit increased errors. This discrepancy highlights the challenges faced in simulating high-speed aerodynamics accurately. The research underscores the importance of improving model fidelity in diverse Mach regimes, particularly in supersonic conditions where traditional methods struggle. Future research directions identified include the development of unsteady flow field models, which are crucial for dynamic analyses, optimization of grid structures for three-dimensional complex fields to improve computational efficiency, and the creation of extensive model libraries. These advancements aim to enhance the accuracy, reliability, and practicality of flow field simulations, extending their applicability in both academic studies and industry applications, particularly in aerospace engineering where precise flow modeling is critical.

Aero-optical effects, part II. Sources of aberrations: tutorial

This paper serves as part II of a two-part tutorial on “aero-optical effects.” In part I, we provide introductory material with an emphasis on system-level considerations, particularly for those who are new to the field of aero-optics. In part II, we move on to survey several sources of aberrations. For example, we cover foundational sources like boundary layers and shear layers, as well as miscellaneous sources like mechanical contamination, shock waves, and aero acoustics. Throughout part II, we emphasize drivers for system-level performance, which appropriately builds on the system-level considerations covered in part I. This emphasis will inform future efforts looking to develop airborne-laser systems flying at subsonic, supersonic, and hypersonic speeds.

Reaction zone structure of spontaneous reaction wave at steady subsonic and supersonic speeds

The reaction zone structure of spontaneous reaction wave at steady subsonic and supersonic speeds is discussed using numerical simulation and theoretical approaches. The main objective of this study is to develop the mathematical model of the spontaneous reaction wave propagating at a steady speed and to clarify the relationship between the reaction zone structure and non-dimensional parameters. A one-dimensional direct numerical simulation (DNS) of the spontaneous reaction wave propagation is performed, and the modeling assumptions are discussed based on the numerical results. The non-dimensional parameters, the Damköhler number and a parameter related to compressibility defined by the Mach number of the spontaneous reaction wave velocity, are introduced into the modeling. The mathematical model is developed using asymptotic techniques, and the solution of the reaction zone structure is analytically derived using Newton’s method. The relationship between reaction zone structure of spontaneous reaction wave and non-dimensional parameters is discussed based on the analytical solutions.

Numerical simulation and experimental study on guidance performance of hypersonic seeker under aerodynamic optical transmission effects

When a hypersonic seeker flies at high speed within the atmosphere, intense interaction with the incoming flow gradually develops into a complex turbulent flow field. This interaction results in complex thermal responses at the seeker window, causing aerodynamic optical effects such as image shift, jitter, and blur of the target image, thereby restricting the seeker's detection capability and accuracy. This paper uses a numerical simulation model for the guidance performance of a hypersonic seeker under aerodynamic optical transmission effects. The study focuses on an ellipsoidal seeker, with its supersonic flight simulation on the basis of the Reynolds-Averaged Navier-Stokes (RANS) equations to get a non-uniform gradient flow field. The correctness of the flow filed results can be verified by wind tunnel experiments. The transient temperature field of the seeker is solved using an unsteady thermal conduction-radiation coupled fluid-solid heat transfer method. Finally, the guidance performance of the hypersonic seeker under aerodynamic optical effects is predicted using the ray tracing method, which employs wavefront aberration, point spread function, degraded images, and image shift.

ВИЗНАЧЕННЯ АЕРОДИНАМІЧНИХ ХАРАКТЕРИСТИК ЛІТАЛЬНОГО АПАРАТА З НАДЗВУКОВИМ ШВИДКОСТЯМИ ПОЛЬОТУ

The aircraft aerodynamic characteristics at supersonic flight modes can be determined with a high level of accuracy through testing the aircraft in wind tunnels or through time-consuming three-dimensional and analytical computational methods. However, in certain cases when the aircraft aerodynamic characteristics are already known within a given Mach number range for supersonic flight, it is possible to develop a relatively simple mathematical model for the aircraft aerodynamic characteristics at supersonic speeds using the aerodynamic properties of the supersonic wing leading edge for flight speeds Mf ≥ 1.2M. The object of the study is the aircraft aerodynamic characteristics at supersonic flight speeds. The subject of the study is the impact of the aircraft geometric features on the mathematical model, which is used to determine its aerodynamic characteristics in the supersonic flight speed range. The study hypothesizes that to determine the aerodynamic characteristics of an aircraft at Mf ≥ 1.2, it is sufficient to identify the values of the aircraft aerodynamic parameters at a flight speed of Mf =1.2 and then use aerodynamic analytical dependencies for the supersonic wing leading edge to determine these parameters at Mf ≥ 1.2. A mathematical model of the aircraft aerodynamic characteristics in the supersonic flight speed range was developed. Coefficients of drag polar and zero-lift drag for the reference mode Mf = 1.2 have been identified by the developed mathematical model. Then, the Mach number range for the aerodynamic characteristics of the NASA 1044 configuration has been extended to supersonic speeds of Mf = 1.8…4.0. The resultsof the aerodynamic calculations, obtained in the form of dependencies of the drag polar and zero-lift drag coefficients on Mach number, and the aircraft polar plot, have been compared with known NASA data for the 1044 configuration aircraft. The average modeling error of the aircraft aerodynamic characteristics in the supersonic speed range is less than 3%.

Impact of surrounding environment on impinging supersonic jets resonances and their hysteretical behavior

Supersonic jets impinging on flat surface produce intense acoustic tones. Recent measurement campaigns have indicated that the surrounding environment geometry or its acoustic treatment can drastically change the observed resonance frequencies and amplitudes at identical flow regimes. For example, at supersonic speeds, a hard walled external nozzle shape leads to the existence of two main types of resonances: antisymmetric resonances (with frequencies close to the free jet screech) and axisymmetric resonances. Frequencies associated to those different resonances are very distinct. We have observed that Applying a simple non-reflective material to the outer surfaces of the nozzle suppresses the axisymmetric resonances and that antisymmetric resonances seems to be contained in the frequency range of existence of trapped acoustics modes, known to be responsible for screech resonances. Additionally, when no acoustic treatment is applied on the nozzle wall, the experiments show that the switch between different resonance regimes (symmetric - antisymmetric), induced by changes in Mach number, has an hysteretical behavior. To the authors' knowledge, this has been rarely discussed in the literature and adds to the actual complexity of the aeroacoustic behavior of such jet flows, especially when modeling is considered.

Energy and exergy analyses on the high-pressure bleeding-air variable cycle designed for a wide-speed-range airbreathing propulsion system

The performance of the traditional turbojet decreases sharply with the increase of the Mach number in the supersonic flight, especially for the turbine-based combined cycle (TBCC) engine in the transition region of the turbojet and scramjet engine. This paper proposes a new high-pressure bleeding-air variable cycle engine (HB-VCE) concept that aims to ease the sharp contradiction between thrust output and fuel consumption of TBCC in the transition region. Through the energy and exergy analyses, the HB-VCE shows greater potential in energy utilization under high Mach flight conditions. The performance of HB-VCE significantly improves as the bleed ratio increases, with a more pronounced effect at higher Mach numbers. At Mach 1.5, the HB-VCE increases the specific thrust (ST) by 2.84 % and reduces the specific fuel consumption (SFC) by 2.73 %. At Mach 2.4, it increases the ST by 6.60 % and reduces the SFC by 6.52 %. Additionally, the controlling law of the critical bleeding ratio constrained by the turbine outlet temperature is proposed. Finally, the effect of the turbine inlet temperature, overall pressure ratio, and Mach number on the critical bleeding ratio is analyzed. This paper provides a promising HB-VCE concept with greater energy utilization for the TBCC design in the transition region.

Aurorally Driven Supersonic Gravity Waves in Saturn's Atmosphere

AbstractSimulations with the Saturn Thermosphere‐Ionosphere General Circulation Model have revealed global‐scale gravity waves in Saturn's upper atmosphere that have not been observed or predicted before. They are forced by diurnally varying zonally non‐uniform distributions of Joule heating and ion drag at the high‐latitude auroral regions in both hemispheres and propagate outward from the source region. The supersonic zonal phase speed imposed by Saturn's rotation and the subsonic meridional phase velocity of about 800 m form a distinct spiral wave structure. They exhibit features of gravity waves: vertical phase progression opposite to the propagation of wave energy and long vertical wavelengths consistent with the dispersion relation for gravity waves. The amplitudes of the perturbations grow with height, reaching 6%–10% for relative temperature variations and up to 350 m for the meridional velocity perturbations. The main effect of these waves is to accelerate the retrograde westerly jets.

Optimization of Plenum for Control of Boundary Layer-Shock Interaction in Supersonic Inlet

In the developing world, the importance of supersonic flights is increasing day by day, especially in military fields. Since supersonic flights have different physical conditions, different designs are required both in the fuselage structure and in the engine part compared to subsonic aircraft. In these studies, the air inlet performance of jet engines designed for supersonic flights is investigated. A plenum has been added to the air inlet geometry in line with the goal of low flow distortion and high pressure recovery. Then, the optimum plenum geometry design was obtained by applying a two-stage optimization process. In each optimization step, computational fluid dynamics analyses were performed to define the effects of the changes in the geometric dimensions of the plenum design and bleed system on the PR (Pressure Recovery) and FD (Flow Distortion) values. The outcome of the analysis showed that the addition of the plenum and the bleed system improved the performance of the air inlet. The new plenum design that emerged as a result of the optimization processes has positively affected the performance values of the air intake. Analysis results showed higher PR and lower FD results in optimized geometry.

Drag Reduction on the Basis of the Area Rule of the Small-Scale Supersonic Flight Experiment Vehicle Being Developed at Muroran Institute of Technology (Second Report)

Drag Reduction on the Basis of the Area Rule of the Small-Scale Supersonic Flight Experiment Vehicle Being Developed at Muroran Institute of Technology (Second Report)

Thermodynamic analysis of the hydrocarbon-fuelled air-turborocket engine with complete-combustion gas generator

The Air-Turborocket (ATR) engine can work at Mach 0∼4 or even higher speed, which is considered one of the best low Mach number propulsion systems for reusable hypersonic vehicles. However, because the hydrocarbon-fuelled ATR engine uses a fuel-rich gas generator, the combustion product contains a large amount of C(gr) that can cause coking in the turbine in a few minutes. To solve this problem, an ATR engine cycle with a complete-combustion gas generator (ATR-CCGG) was proposed. The performance of this cycle has been analysed through the thermodynamic model, and the influence factors and the sensitivity study of the cycle performance have been investigated. The results show that when the equivalence ratio is 1, the cycle can get more than 700 s of specific impulse and 1000 m/s of specific thrust at supersonic speed. Although the performance at subsonic speed is lower than that of the LOX/Kerosene ATR engine, the gas generator without C(gr) can ensure the engine to work for hours without coking in the turbine at different Mach numbers, which can be used in reusable hypersonic vehicles or single-stage-to-orbit missions.

Aerodynamic Drag Coefficient Prediction of a Spike Blunt Body Based on K-Nearest Neighbors

Spike blunt bodies are a method to reduce drag when a body moves at speeds above sound. Several numerical works based on computational fluid dynamics (CFD) have deeply studied fluid performance and highlighted its advantages. However, most documentation focuses on modifying spike physical properties while keeping constant supersonic or hypersonic flow conditions. In recent years, machine learning models have emerged as viable tools to predict values in almost any field, including aerodynamics. In the case of CFD, many models have been explored, such as support vector regression, ensemble methods, and artificial neural networks. However, a simple and easy-to-implement method such as k-Nearest Neighbors has not been extensively explored. This work extrapoled k-Nearest Neighbors to predict the drag coefficient of a spike blunt body for a range of supersonic and hypersonic speeds based on drag data obtained from CFD analysis. The parametric study of the spike blunt body was performed considering body diameter, spike length, and freestream Mach number as input variables. The algorithm presents proper predictions, with errors less than 5% for the drag coefficient and considering a minimum of three neighbor nodes. The k-NN was compared again Kriging model and k-NN presents a better accuracy. The above validates the flexibility of the method and shows a new area of opportunity for the calculation of aerodynamic properties.

Mixed-Flow Turbofan Engine Model for the Conceptual Design of Sustainable Supersonic Airplanes

Current research efforts on commercial supersonic flight aim to overcome past challenges by designing a new generation of sustainable supersonic airplanes. Achieving this goal requires careful consideration of the propulsion system during the design process. This study proposes a mixed-flow turbofan engine model coupled with emission estimation routines to increase the reliability of the conceptual design of future supersonic aircraft. The model enables parametric analyses by analyzing variations in main engine design parameters (πc,πf, BPR) as function of the system and mission requirements, such as the Mach number, and suggesting applicability boundaries. The overall methodology was applied to a low-boom Mach 1.5 case study, allowing for both on-design and off-design analyses and generating a propulsive database to support preliminary mission simulations and chemical emission estimation. Finally, the accuracy and reliability of the engine model was validated against GSP 11 data for a generic mixed-flow turbofan engine. A modified version of the Fuel Flow Method, originally developed by Boeing, allowing for emissions estimation throughout the mission for a supersonic engine using biofuels. The application of the methodology led to the definition of an engine with a πc of 30 and BPR of 0.7 for the selected case study, which was successful in meeting the initial mission requirements.

A Practical Approach for Estimating Drag Increase due to Volume Increase in the Center Body of a High-Speed UAV

High-speed unmanned aerial vehicles are used for many purposes in aviation. High-speed aircraft do not only fly at supersonic speed, but their subsonic flight performance is also important. Aircraft design is an iterative process in which many disciplines work together. The design process is updated by the negotiation of different disciplines. There may be a demand for a body volume increase in the interior design process. The increase in volume can be achieved by the elongation or expansion of the body. A volume increase in the center body causes an increase in drag. The aim of this study is to predict the effect of the elongation or expansion of the center body on drag in a practical way. It is investigated using a structure proposed with MATLAB and DATCOM. The results in both subsonic and supersonic regimes are formalized separately and compared. It is shown that in case of providing the same volume increase in both subsonic and supersonic regimes, the elongation of the aircraft center body causes less drag compared to widening.

Innovative Metal Powder Production Using CFD with Convergent-Divergent Nozzles in Wire Arc Atomization

This study aims to enhance the production of metal powders using a novel approach that integrates computational fluid dynamics (CFD) with convergent-divergent (C-D) nozzles in wire arc spraying atomization (WASA). The primary objective is to investigate the influence of nozzle design on particle size distribution and production efficiency. Utilizing the ANSYS CFD Fluent program, simulations were conducted to analyze the effects of various parameters, including throat diameter and divergent angles, on gas dynamics and metal droplet behavior. The findings reveal that C-D nozzles facilitate the acceleration of gas flow to supersonic speeds, significantly improving the shear force acting on the molten metal, thereby promoting the fragmentation of droplets into smaller particles. Notably, the optimized nozzle configuration achieved a median particle size (D50) of 44.42 µm, suitable for additive manufacturing applications. The novelty of this work lies in its comprehensive simulation framework that allows for rapid virtual testing, potentially leading to significant improvements in the efficiency and quality of metal powder production processes. This research addresses critical gaps in the existing literature and provides a robust foundation for future studies in the field of metal powder manufacturing. Doi: 10.28991/HIJ-2024-05-03-02 Full Text: PDF