Supersonic Aircraft Research Articles

Experimental Characterization of Supersonic Diamond Jet Acoustics

Acoustic emissions from supersonic jets are of interest due to their deleterious effects on nearby personnel and structures. As development of both military and commercial supersonic aircraft continues to grow, it is essential to combat increasing noise pollution. Axisymmetric jets are highly prevalent where their aeroacoustic characteristics have been well-documented; in contrast, relatively few studies have examined non-axisymmetric jets. In the present study, aeroacoustics of supersonic jets produced by a diamond-shaped Mach 2 nozzle are explored across a large operational envelope under different pressure ratios ranging from highly overexpanded to moderately underexpanded. Measurements of the far-field acoustics in an anechoic environment were used to characterize the acoustic flowfield and contrast the distribution of energy in the spectral domain between the characteristic major and minor axes of the nozzle. Overall, the diamond jet exhibits a highly directional acoustic field with a distinct local minimum for sideline noise. Furthermore, screech behavior was only observed in a narrow, overexpanded flow regime; this is in contrast to the documented behavior of axisymmetric and rectangular jets. Over the conditions examined there is no indication of distinct staging. Additionally, there exists a nozzle pressure ratio threshold beyond which peak noise switches axes.

Low-Boom Design for Supersonic Transport with Canard and Forward-Swept Wings Using Equivalent Area Design Method

Forward-swept wings can be expected to be lower-boom planforms with similar amount of drag as backward-swept wings because of their good lift distributions. In this study, the equivalent area distribution of a ten-seater supersonic forward-swept wing aircraft with a canard was designed to obtain design knowledge for leading boom reduction. The equivalent area distribution of the aircraft was calculated by solving the compressible Euler equation. A feasible target equivalent area distribution was generated based on Darden’s method and compared with the equivalent area distribution. To achieve a closer match in terms of lift and geometry with the target, the main wing planform and the position of the main wing along the body and vertical axes were modified. The low-boom performances were evaluated using the extended Burgers equation. The design results indicated that the forward-swept wing configuration with a canard could divide the single peak of the leading boom into two peaks. Thus, the sonic boom strength of the canard configuration was 2.5 PLdB lower than that of the configuration without the canard wing.

Direct Method of Aerodynamic Shape Optimization For Supersonic Aircraft Design

The efficient direct optimization method for Supersonic Aircraft Design is presented. Aerodynamic drag and sonic boom signature parameters are used as the objective functions minimized under geometric and aerodynamic constraints. The method combines Newton based algorithm with Euler space-marching solver. The efficiency of the method is demonstrated on examples of two- and three-dimensional aerodynamic design. Nose shapes that ensure minimum wave drag for specified constraints on the overall size are defined. It is shown that the near-optimal axisymmetric noses have a flat forward face. The median surface of a complex plane form wing with minimal drag due to lift is constructed. The simplest shape deformation of the wing providing the main part of aerodynamic drag diminishing is established. To profile the aircraft wing and fuselage providing optimal sonic boom parameters approaches of theory for sonic boom propagation are included in the method. The initial shock intensity as the primary value of the pressure signature is minimized. The results obtained within the framework of the Eulers equations and simplified flow models are compared.

Methods for determining the weight of a supersonic passenger administrative aircraft fuselage structure based on regression analysis

Methods for determining the weight of a supersonic passenger administrative aircraft fuselage structure based on regression analysis

A Rapid Surrogate Model for Estimating Aviation Noise Impact across Various Departure Profiles and Operating Conditions

Aviation noise remains a key barrier to the sustainable growth of commercial aviation. The advent of emerging technologies, such as urban air mobility, and the renewed interest in commercial supersonic transport aircraft, has only further raised concerns over the resultant community noise exposure. The foundation of any noise mitigation effort is the ability to accurately model noise metrics over a wide range of scenarios. Aviation noise is influenced by a wide variety of factors, including aircraft type, payload weight, thrust settings, airport elevation, ambient weather, and flight trajectory. Traditional noise modeling paradigms rely on physics-based and empirical calculations, which are computationally expensive. Attempts at speeding up the computations with alternate models could deliver on speed or accuracy, but not both. Recent research has indicated that model order reduction techniques hold promise for transforming and greatly reducing the number of quantities that need to be modeled. Paired with surrogate modeling techniques, a rapid and accurate noise model can be generated. The research presented in this manuscript expands on the model order reduction method and develops a rapid noise surrogate model, which can account for the piloting actions, the ambient temperature, and airport elevation. The presented results indicate that the method works well with minimal error for most modeling scenarios. The results also outline avenues for improvement, such as using a different class of surrogate models or modeling additional training cases. The model developed in this research has numerous applications for multi-query applications, such as parametric trade-off analyses and optimization studies. With the inclusion of airport and aircraft parameters, the model enables the development of frameworks that optimize piloting actions for noise mitigation on the ground.

Survey of Semi-Empirical Jet Noise Models for Preliminary Aircraft Engine Design

Scientific research studies on jet noise generation have been ongoing since the early 1950s, when turbojets were first used in commercial aircraft. Several numerical methods have been developed with the aim of reducing the environmental issues related to the impact of jet noise on community annoyance. Among them, the development of fast and comprehensive tools for jet noise prediction captured the attention of researchers and engineers, being very useful in the preliminary design phase of aircraft engines. This work deals with an extensive survey of James R. Stone’s models, initially formulated for the National Aeronautics and Space Administration (NASA) by Modern Technologies Corporation (MTC), and their implementation in a contemporary numerical framework. The models and their implementation are validated by simulating different engine settings and nozzle configurations taken from the literature with the main scope of highlighting the strengths and weaknesses of the different semi-empirical formulations and setting guidelines for their effective use during the design phases of the next generation of supersonic aircraft.

Constraints in the Problem of Calculating Optimal Trajectories for a Supersonic Non-Maneuverable Aircraft

The influence of phase and other constraints on the method of searching for the trajectories of the movement of a civil supersonic aircraft, which are optimal in terms of fuel consumption, is considered. Based on the solutions found by the dynamic programming method, taking into account numerous restrictions on flight altitude, pitch angle, normal high-speed overload, aircraft speed, engine thrust, etc., it is shown that almost all of these conditions can be ignored during the initial stage of calculations, since the optimal solution does not reach them. Therefore, one can first apply the maximum principle, and use the dynamic programming method only in those cases where a substantial part of the constraints turns out to be significant.

Conditions for simulation of transonic flutter of aerodynamic control surfaces of supersonic aircraft in wind tunnels

The transonic flutter of the aerodynamic control surfaces of supersonic and hypersonic aircraft refers to those phenomena of dynamic aeroelasticity, the assessment of whichin a flight experiment is dangerous. Since there is still no universally accepted mathematical model of the occurrence of this phenomenon, tests of dynamic-like models in wind tunnels can be classified as basic and safe methods for evaluating the characteristics of transonic flutter. Therefore, the substantiation of the conditions for modeling transonic flutter, which allow the transfer of the results of blowing of dynamic-like models in wind tunnels to full-scale aircraft designs, remains an actual scientific problem.The article proposes one of the possible approaches to justifying the conditions for modeling transonic flutter of aerodynamic control surfaces in wind tunnels, which is based on the analysis of a nonlinear mathematical model of the occurrence of this phenomenon.Based on the analysis of this mathematical model, it was determined that, in addition to the conditions for modeling transonic flutter in wind tunnels, which are due to the geometric similarity of the system "carrying aerodynamic surface – aerodynamic control surface", additional conditions for modeling this phenomenon should be the following dimensionless quantities of nature and model:- amplitudes of oscillations of nature and model;- numbers M at which transonic flutter occurs;- logarithmic decrements of oscillations of aerodynamic control surfaces;- Strouhal numbers;- ratio of gas density to material density of aerodynamic control surfaces;- the adiabatic index k=1.405, that is, the working body in the wind tunnel should be air.Blowing of dynamically similar models must be carried out in wind tunnels of a continuous type, subject to conditions (s-1).

Effect of Air Jet Vortex Generators on the Shock Wave Boundary Layer Interaction of Transonic Wing

The interaction between shock waves and turbulent boundary layers (SBLI) is a common phenomenon in transonic and supersonic aircraft wings. In this study, we simulated the SBLI of a classical NACA0012 wing at an angle of attack (AOA) of 1.4° and Mach number (Ma) of 0.78 using the open-source software OpenFOAM. Our results show that an air-jet vortex generator can effectively reduce the length of the separation zone and improve the lift coefficient of the airfoil. The vortex structure generated by the jet vortex generator significantly reduces the separation caused by SBLI. We conducted simulations with jet angles of 30°, 45°, and 60° and found that the larger the jet angle, the stronger the vortex and the greater the improvement in the lift coefficient. When the jet angle was 60°, the vortex structure generated by the jet vortex generator transformed the normal shock wave into a λ shock wave, resulting in a maximum increase in the lift coefficient of 2.35%. The simulations focused on exploring the effect of the jet angle and determined that that optimal jet parameters that effectively reduce SBLI damage and improve the lift coefficient of the airfoil.

Working state map of hydrocarbon fuels for regenerative cooling

Regenerative cooling by endothermic hydrocarbon fuel (EHF) is one of the most promising techniques for thermal management of supersonic or hypersonic aircraft. How to maintain the fuel working in proper states is an important issue to maximize the cooling potential of EHT. This work proposes a novel working state map, including risking zone (RZ), thermal cracking zone (TCZ), supercritical zone (SupZ) and subcritical zone (SubZ), to differentiate possible working states of an EHF during regenerative cooling. Using n-decane flowing in a circular tube as an example, the boundaries of four zones are determined by numerical simulation covering different heat fluxes (0.2–4.0 MW·m−2) and mass flow rates (0.5–10.5 g·s−1) under two operating pressures (3.45 and 5.00 MPa). Empirical correlations for three boundary lines are obtained and the maximum cooling capacity is identified, as well as the identification of the pressure effect. The revelation of such new perspective of regenerative cooling is of great implication to the design and optimization of cooling system for future thermal management.

Preliminary design of next generation Mach 1.6 supersonic business jets to investigate landing & take-off (LTO) noise and emissions – SENECA.

With the approach of next generation supersonic transport entry into service, new research activities were initiated to support updates on ICAO regulations and certification processes for supersonic transport vehicles. Within this context, the EU Horizon 2020 SENECA project has been launched to investigate the levels of noise and gaseous emissions in the vicinity of airports as well as the global climate impact of next generation supersonic civil aircraft. This paper introduces some of the preliminary outcomes of this investigation. It presents the preliminary design and performance analysis of a Mach 1.6 business jet, following an integrated aircraft-engine design approach. The preliminary design was performed accounting for the limitations posed by future environmental restrictions on respective subsonic vehicles. The market space and mission route definition exercise assumed only “over-sea” supersonic operations, while for “over-land”, only subsonic operations where allowed. Parametric studies on engine integrated design demonstrated modest core temperatures while cruising and the significant impact of engine installation on performance. At this first design iteration, assuming current state of the art technology, the Mach 1.6 business jet showed good potential to satisfy the predicted mission requirements while respecting the environmental constraints in terms of Landing & Take-Off (LTO) noise and emissions.

How quiet is the quietest supersonic aircraft?

Estimations of the noise level of NASA’s experimental X-59 supersonic aircraft will help inform future community flyovers.

European Defence Planning: Concorde, Franco-British Politico-Military Relations and the Cold War, 1956–68

Cold War tensions between the superpowers marked the years 1956 to 1966, where the United States, Britain and France prioritised European defence against Soviet aggression. Despite being envisaged for civil aviation purposes, the Concorde aircraft was the Franco-British alternative to US proposals concerning the introduction of a supersonic bomber within the Inter-Allied Nuclear Force (IANF) to protect against Soviet attack during the Cold War technological race. This offers a unique case study for the examination of Franco-British bi-lateral partnership in the context of European defence. This article will concentrate on three main themes. The first investigates how Concorde was considered as a viable component for the IANF. Subsequently, the loci of US, British and French policy decisions will be explored with regards to constructing supersonic aircrafts. Lastly, the article considers how foreign policies influenced the shift from a military-use Concorde to a more commercial option. The historiography on Concorde focuses on its commercial impact, and how it affected the Franco-British partnership. Considering Concorde from the defence perspective allows us to analyse the divisions between the US and French grand design ideas – the IANF and Europe puissance – and how they provoked further friction in the Franco-British military relationship.

The Study of Selected Aspects of the Suborbital Vehicle Return Flight Trajectory

The article presents the results of preliminary studies of the parameters of the return flight trajectory of a rocket plane for suborbital tourist flights into space. The rocket plane is designed as a tailless vehicle and has an unconventional arrangement of control surfaces: elevons and side plates that can rotate. The main aim of the research presented in this paper is to investigate the dynamic stability of the rocket plane and the response to control in the return suborbital flight. The secondary objective is to study the behavior of the rocket plane with respect to the initial state of the return flight. The key parameters taken into account in this study are the Mach number and G-load. Moreover, a study of the trim condition, dynamic stability and response to control of a rocket plane in the low part of the stratosphere is presented. The tests were carried out using a numerical simulation of the flight of a rocket plane. Dynamic stability was determined on the basis of time history analysis, and the results were compared with the results obtained by solving the eigenvalues problem. The results revealed that the rocket plane should be equipped with a Stability Augmentation System to improve short period damping at supersonic speeds at moderate altitudes. It can also be concluded that the maximum load G and Ma do not occur at the same height of flight. In terms of the effectiveness of the control surfaces, they start working at an altitude of 55 km. Due to the speed regime, the obtained results can be useful in the design of such objects as rocket planes, highly maneuverable and supersonic aircraft.

The problems of creating a propulsion system of a new generation supersonic passenger aircraft (review)

The problems of creating a propulsion system for a new generation supersonic passenger aircraft are considered on the basis of a review of the work on the supersonic transport being carried out in the world. It is shown that the desire to achieve high flight performance and commercial effectiveness of a supersonic passenger aircraft while meeting up-to-date environmental requirements leads to contradictory technical solutions regarding the propulsion system: the location and number of engines, the scheme of the air intake and nozzle, the choice of the scheme and design parameters of the engine, the use of new high-temperature materials in the engine hot section, etc. The features of the operating conditions of the engine components of a supersonic passenger aircraft in comparison with the engines of up-to-date subsonic civil aviation aircraft and supersonic military aircraft are indicated. The calculated estimates of the influence of various technical solutions on the parameters of the supersonic passenger aircraft engine are given. Due to the complexity and multi-criterion nature of the task of creating a supersonic passenger aircraft propulsion system, its solution requires an integrated approach based on close cooperation of specialists in airframe aerodynamics, engine, etc.

Aerodynamic Optimization Design of Supersonic Wing Based on Discrete Adjoint

Reducing fuel consumption and improving the economy by effectively reducing cruising drag is the main objective of the aerodynamic design of supersonic civil aircraft. In this paper, the aerodynamic optimization design system based on the Reynolds-Averaged Navier–Stokes (RANS) equation and discrete adjoint theory is applied to supersonic wing design. Based on this system, a single-point optimization design study of aerodynamic drag reduction in cruise conditions was carried out for two typical supersonic wing layouts, subsonic leading edge and supersonic leading edge, and the drag reduction reached 3.78% and 4.53%, respectively. The aerodynamic design characteristics of different types of supersonic wings were explored from the perspectives of wing load, twist angle distribution, pressure distribution, airfoil shape characteristics, and flow field characteristics. The optimization results show that the drag reduction of the subsonic leading edge configuration is dominated by the induced drag, while the optimizer mainly focuses on reducing the shock wave drag for the supersonic leading edge configuration. By comparing the sensitivity analysis of lift and drag coefficients to airfoil deformation with the optimization results, the optimized dominant directions of two types of supersonic wings are qualitatively analyzed. The derivatives obtained from discrete adjoint equations are useful to elaborate the design tendency and the reason for the trade-off generation of supersonic wings under specific layouts and engineering constraints, which provides a reference for the design of supersonic wings in the future.

Review of vortex lattice method for supersonic aircraft design

AbstractThere has been a renewed interest in developing environmentally friendly, economically viable, and technologically feasible supersonic transport aircraft and reduced order modeling methods can play an important contribution in accelerating the design process of these future aircraft. This paper reviews the use of the vortex lattice method (VLM) in modeling the general aerodynamics of subsonic and supersonic aircraft. The historical overview of the vortex lattice method is reviewed which indicates the use of this method for over a century for development and advancements in the aerodynamic analysis of subsonic and supersonic aircraft. The preference of VLM over other potential flow-solvers is because of its low order highly efficient computational analysis which is quick and efficient. Developments in VLM covering steady, unsteady state, linear and non-linear aerodynamic characteristics for different wing planform for the purpose of several different types of design optimisation is reviewed. For over a decade classical vortex lattice method has been used for multi-objective optimisation studies for commercial aircraft and unmanned aerial vehicle’s aerodynamic performance optimisation. VLM was one of the major potential flow solvers for studying the aerodynamic and aeroelastic characteristics of many wings and aircraft for NASA’s supersonic transport mission (SST). VLM is a preferred means for solving large numbers of computational design parameters in less time, more efficiently, and cheaper when compared to conventional CFD analysis which lends itself more to detailed study and solving the more challenging configuration and aerodynamic features of civil supersonic transport.

Potential Impacts on Ozone and Climate From a Proposed Fleet of Supersonic Aircraft

AbstractThere has been renewed interest in developing commercial supersonic transport aircraft due to the increased overall demands by the public for air travel, the aspiration for more intercontinental travel, and the desire for shorter flight times. Various companies and academic institutions have been actively considering the designs of such supersonic aircraft. As these new designs are developed, the environmental impact on ozone and climate of these fleets need to be explored. This study examines one such proposed commercial supersonic fleet of 55‐seater that is projected to fly at Mach 2.2, corresponding to cruise altitudes of 17–20 km, and which would burn 122.32 Tg of fuel and emit 1.78 Tg of NOx each year. Our analyses indicate this proposed fleet would cause a 0.74% reduction in global column ozone (∼2 Dobson Units), which is mainly attributed to the large amounts of nitrogen oxides released in the atmosphere from the supersonic aircraft. The maximum ozone loss occurs at the tropics in the fall season, with a reduction of −1.4% in the total column ozone regionally. The stratospheric‐adjusted radiative forcing on climate from this fleet was derived based on changes in atmospheric concentrations of ozone (59.5 mW/m2), water vapor (10.1 mW/m2), black carbon (−3.9 mW/m2) and sulfate aerosols (−20.3 mW/m2), resulting in a net non‐CO2, non‐contrail forcing of 45.4 mW/m2, indicating an overall warming effect.

Effect of longitudinal lift distribution on sonic boom of a canard-wing-stabilator-body configuration

After the last flight of the Concorde in 2003, sonic boom has been one of the obstacles to the return of a supersonic transport aircraft to service. To reduce the sonic boom intensity to an acceptable level, it is of great significance to study the effect of lift distribution on far-field sonic boom, since lift is one of the most important contributors to an intense sonic boom. Existing studies on the longitudinal lift distribution used low-fidelity methods, such as Whitham theory, and in turn, only preliminary conclusions were obtained, such as that extending the lift distribution can reduce sonic boom. This paper uses a newly developed high-fidelity prediction method to quantitatively study the effect of longitudinal lift distribution on the sonic boom of a Canard-Wing-Stabilator-Body (CWSB) configuration. This high-fidelity prediction method combines near-field CFD simulation with far-field propagation by solving the augmented Burgers equation. A multipole analysis method is employed for the extraction of near-field waveform in order to reduce computational cost. Seven configurations with the same total lift but different distributions are studied, and the quantitative relationship between the longitudinal lift distribution and far-field sonic boom intensity is investigated. It is observed that a small lift generated by the stabilator can prevent aft-stabilator and aft-fuselage shocks from merging, while the balanced lift generated by the canard and wing can effectively keep the corresponding shocks further apart, which is beneficial for reducing both the on-track and off-track sonic boom. In turn, the acoustic level perceived at the ground can be reduced by 5.9 PLdB on-track and 5.4 PLdB off-track, on average.

Design and Verification of Short-Distance Landing Control System for a One-Third-Scale Unmanned Supersonic Experimental Airplane

The Aerospace Plane Research Center at the Muroran Institute of Technology is currently conducting research to develop enabling technologies for high-speed aircraft traveling at high altitudes and constructing experimental, small-scale, unmanned supersonic aircraft called Oowashi as a testbed for flight. To confirm the control performance of the aircraft, an experiment using a one-third-scale model of the Oowashi aircraft has been planned. The flight of high-speed aircraft always presents the problem of having to land on an ordinary runway regardless of the aircraft’s high speed at the beginning of the landing process. This paper therefore proposes a new landing control design method that can shorten the landing distance for a high-speed aircraft without increasing the rate of descent. The design method utilizes the newly clarified relationship between an angle of attack and the time constant of flare control system, which is effective to raise glideslope angle during landing. The validity of the method is confirmed by computer simulation assuming the model aircraft equivalent to a one-third-scale model of the Oowashi aircraft.