Cruise Mach Number Research Articles

Numerical Study of Variable Camber Continuous Trailing Edge Flap at Off-Design Conditions

Numerical simulations are performed to study the outboard airfoil of advanced technology regional aircraft (ATRA) wings with five different variable camber continuous trailing edge flap (VCCTEF) configurations. The computational study aims to improve the aerodynamic efficiency of the airfoil under cruise conditions. The design of outboard airfoil complies with the hybrid laminar flow control design criteria. This work is unique in terms of analysis of the effects of VCCTEF on the ATRA wing’s outboard airfoil during the off-design condition. The Reynolds–Averaged Navier–Stokes equations coupled with the Spalart-Allmaras turbulence model are employed to perform the simulations for the baseline case and VCCTEF configurations. The current computational study is performed at an altitude of 10 km with a cruise Mach number of 0.77 and a Reynolds number of 2.16 × 107. Amongst all five configurations of VCCTEF airfoils studied, a flap having a parabolic profile (VCCTEF 123) configuration shows the maximum airfoil efficiency and resulted in an increase of 6.3% as compared to the baseline airfoil.

Robust optimization for a wing at drag divergence Mach number based on an improved PSO algorithm

In this study, an improved PSO (particle swarm optimization) algorithm is proposed and applied to the robust optimization of a wing at drag divergence Mach number. In order to reduce the number of design variables, a six-order CST (class/shape function transformation) method is employed for airfoil parameterization. For the purpose of improving the optimization efficiency, Delaunay graph mapping method is adopted for mesh deformation in each iteration of the airfoil optimization, and NURBS (non-uniform rational B-splines)-FFD (free-form deformation) method is employed for mesh deformation in each iteration of the wing optimization. For improving the standard PSO algorithm, CVTs (centroidal Voronoi tessellations) method is introduced to generate original positions of the particles more dispersedly, a second-order oscillating scheme is used and an FDR (fitness distance ratio) item is added for updating velocities and positions of the particles. By virtue of the improved PSO algorithm, single point optimization and robust optimization are conducted for both airfoil and wing. The results indicate that, comparing with the single point optimizations, the robust optimizations not only reduce drag coefficients of the airfoil and the wing at cruise Mach numbers, but also attenuate the drag increments as the Mach number increases up to drag divergence Mach numbers.

Visualization of streamline tracing inlet-isolator flows using a planar laser Rayleigh scattering imaging technique

Isolator flows downstream of the scoop model inlet were visualized by using a planar laser Rayleigh scattering imaging technique. The scoop model was designed for a cruise Mach number of 6. The isolator flows were visualized for two contraction ratios of the scoop inlet (4 and 5) and various freestream Reynolds number conditions. Two freestream Mach numbers were tested to study the flows at the on-design (Mach number 6) and off-design (Mach number 4.5) conditions. Pseudo-cross-sectional flow structures of the isolator were constructed by using averaged planar images from multiple planar imaging planes. The flow visualization found that there exist curved shockwaves and localized flow separation, and three-dimensional shockwave–boundary layer interaction occurs locally affecting boundary layer growth. The visualization showed that Reynolds number determines the thickness of the boundary layer and the size of eddies. Similar overall flow and shockwave structures were observed for different freestream Mach number and Reynolds number conditions. Differences in shockwave angles, impinging shock location, the size of the core flow, and the size of the flow structure were discussed for various flow conditions.

Aerodynamic Shape Optimization of Natural-Laminar-Flow Wing Using Surrogate-Based Approach

Aerodynamic shape optimization of a swept natural-laminar-flow wing in the transonic regime is still challenging. The difficulty is associated with reliable prediction of laminar–turbulence transition and reasonable compromise of viscous and wave drags. This paper proposes to use efficient global optimization based on surrogate models to address this problem. The Reynolds-averaged Navier–Stokes flow solver features automatic transition prediction via a full method, in which dual factors are used for Tollmien–Schlichting and crossflow instabilities, respectively. The optimizer is based on the kriging surrogate model and parallel infill-sampling method. The baseline natural-laminar-flow wing for short- and medium-range transport aircraft is designed at a cruise Mach number 0.75. Then, drag minimization with up to 42 design variables is carried out, and significant drag reduction (8.79%) has been achieved. A close examination of the optimal wing shows that the drag reduction mainly comes from shock-wave weakening on the upper surface and laminar flow extending via suppression of crossflow instability on the lower surface. Robustness of the optimal wing is investigated, and multipoint optimization is further exercised to improve the robustness to the Mach number variation. It is demonstrated that surrogate-based optimization is feasible and effective for aerodynamic shape optimization of transonic natural-laminar-flow wings.

Experimental and numerical investigation of flight dynamics of a generic lambda wing configuration

The development and comparison of aerodynamic simulation models for the inclusion in an open-loop stability & control model based on both experimental and numerical data are presented. Specifically, the work is focused on the effectiveness of the Kestrel simulation tools in replicating the aerodynamic coefficient data from wind tunnel experiments of a generic lambda wing configuration with four trailing edge control surfaces. Simulation conditions correspond to the angles of attack in range of −5∘ to 10∘ at assumed cruise Mach number of 0.7. These simulations include a clean configuration with no control surface deflection and when inboard/outboard control surfaces are deflected −10∘, −5∘, 5∘, and 10∘. Single and overset grids are used to simulate the control surface deflections. Predicted aerodynamic data are collected in form of look-up tables and are run in a Matlab/Simulink program to determine stability and control properties of the vehicle. The vehicle stability and control characteristics are then compared with the results of running experimental coefficient values through the same program. The results show that numerical predictions of the test case considered depend heavily on the geometry details and numerical setup, e.g. turbulence modeling, grid refinement, integration scheme, time step, etc. and therefore do not match experimental data at all tested conditions. In addition, the stability and control investigation of the vehicle using these data confirms that the differences between simulations and measurements could lead to significant changes in the evaluation of vehicle's stability and control characteristics. The simulation predictions of this work, therefore, cannot fully replace the need for experiments to determine the vehicle's stability and control, however they help to provide insight into flow physics and to direct/make best use of wind tunnel experiments.

Cruise Range in Formation Flight

A new set of analytical range equations (a modification of the traditional Bréguet range equation) suitable for formation flight at transonic flight speeds under realistic operating conditions (constant Mach number and altitude) is derived. Formations of two aircraft of the same type are analyzed to determine the effects of 1) weight differences between the aircraft, 2) altitude, and 3) the formation flight range on the potential fuel benefits and the associated optimum Mach number. In the case of a weight difference, the lightest aircraft should lead the formation to realize the largest fuel benefits. Overall, fuel savings of 6 to 12% for the total formation can be realized at the expense of a reduction in cruise Mach number from 0.85 to 0.80. The fuel benefit is much less (2 to 8%) when the formation is flown at the original design cruise Mach number. In terms of fuel benefits and Mach number, it is beneficial to fly in formation at higher altitudes. Formation flight step climb procedures are possible, but the additional fuel savings are minimal compared to a constant-altitude formation flight.

Graphical exergy analysis for a scramjet thermodynamic performance evaluation

Graphical exergy analysis was applied to a scramjet at cruise conditions. A zero-dimensional scramjet model was developed to describe the thermodynamic process of a hypersonic propulsion system. Energy utilisation diagrams (EUDs) were drawn to present the relationship between energy and energy level at cruise conditions. Performance optimisation based on the exergy efficiency was completed ranging from Ma 5.35 to Ma 6.23. The results indicate that the exhausting process produces the most exergy destruction, then the combustion process, and the exergy loss in the expansion process is the smallest. There is a maximal exergy efficiency with the variation of fuel equivalence ratio for a certain entrance condition, and the maximum increases with the cruise Mach number and the altitude. The advantages of graphical exergy analysis, such as visual representation of the relationships among different components and the convenience of analysis, were also shown by the example in this paper.

Assessment of Aircraft Contrail Avoidance Strategies

Some contrail avoidance options are critically analyzed. These options include the use of altitude flexibility and go-around. It is demonstrated that long-range commercial airplanes powered by high bypass turbofan engines have a vertical flexibility of up to 3,000 ft (30 flight levels) for a fuel penalty of 1%. This penalty can be recovered by adjusting the cruise Mach number. However, there are technical difficulties in implementing this option, due to air traffic constraints and safety issues. Avoidance proposals that involve a go-around are impractical, not least for the extension of the flight time, which bears additional operating costs. A sensitivity analysis is shown to indicate that single-event forecasts may be imprecise and that contrail events cannot be costed by precise methods. The paper concludes with the statement that long-term contrail effects (years or decades) obtained by single-flight procedure changes (if allowed at all) cannot be reliably predicted. Simple prediction methods cannot be used for long-term forecasting.

Flow control of a boundary layer ingesting serpentine diffuser via blowing and suction

The use of serpentine boundary layer ingesting (BLI) diffusers offers a significant benefit to the performance of Blended Wing Body (BWB) aircraft. However, the inherent diffuser geometry combined with a thick ingested boundary layer creates strong secondary flows that lead to severe flow distortion at the engine face, increasing the possibility of engine surge. This study investigated the use of enabling active flow control methods to reduce engine-face distortion. An ejector-pump based system of fluidic actuators was used to directly manage the diffuser secondary flows. This system was modeled computationally using a boundary condition jet modeling method, and tested in an ejector-driven wind tunnel facility. This facility is capable of simulating the high-altitude, high subsonic Mach number conditions representative of BWB cruise conditions, specifically a cruise Mach number of 0.85 at an altitude of 12,000 m. The tunnel test section used for this experiment was designed, built, and tested as a validation tool for the computational methods.

Aircraft Route Optimization for Formation Flight

We quantify the fuel and cost benefits of applying extended formation flight to commercial airline operations. Central to this study is the development of a bi-level, mixed-integer real formation flight optimization framework. The framework has two main components: 1) a continuous-domain aircraft mission performance optimization and 2) an integer optimization component that selects the best combination of optimized missions to form a formation flight schedule. The mission performance reflects the effects of rolled-up wakes, formation heterogeneity, and formation-induced compressibility. The results show that an airline can use formation flight to reduce fuel burn by 5.8% or direct operating cost by 2.0% in a long-haul international schedule. The savings increase to 7.7% in fuel or 2.6% in cost for a large-scale, transatlantic airline alliance schedule. These results include the effects of a conservative fuel reserve for formation flight. Sensitivity studies show that a modest reduction in the cruise Mach number may be sufficient to manage the impact of formation-induced compressibility effects on system-level formation flight performance. We demonstrate that the potential savings from extended formation flight (an operational improvement using existing aircraft) can approach those claimed for advanced vehicle technologies and unconventional configurations.

Fifth Drag Prediction Workshop: Computational Fluid Dynamics Studies Carried Out at ONERA

This article presents the computational studies that have been performed at ONERA–The French Aerospace Lab in the context of the 5th AIAA Drag Prediction Workshop. A grid-convergence process and a buffet-onset study have been carried out using a Reynolds-averaged Navier–Stokes solver and far-field methods. In this work, the multiblock grids of the Common Research Model provided by the 5th AIAA Drag Prediction Workshop Committee are employed, and a new meshing strategy involving overset techniques and Cartesian grids is tested. All of these meshes are computed with the ONERA elsA code using the Spalart–Allmaras turbulence model and far-field drag analyses are performed with the ONERA ffd72 software. The grid-convergence study led with the common multiblock grids gives an extrapolated drag value close to 250 drag counts, which corresponds to the numerical median of all 5th AIAA Drag Prediction Workshop participants. The meshing approach using Cartesian grids also provides promising results. Comparisons with NASA wind-tunnel tests show reasonable discrepancies between computational fluid dynamics predictions and experimental drag measurements. The buffet-onset study completed with the medium common multiblock grid is aimed at investigating an angle-of-attack sweep at cruise Mach number. Global and local analyses combining near-field and far-field approaches are presented.

Inverse Design of Low-Boom Supersonic Concepts Using Reversed Equivalent-Area Targets

A promising path for developing a low-boom configuration is a multifidelity approach that starts from a low-fidelity low-boom design, refines the low-fidelity design with computational fluid dynamics equivalent-area analysis, and improves the design with sonic-boom analysis by using computational fluid dynamics off-body pressure distributions. The focus of this paper is on the third step of this approach, in which the design is improved with sonic-boom analysis through the use of computational fluid dynamics calculations. A new inverse design process for off-body pressure tailoring is formulated and demonstrated with a low-boom supersonic configuration that was developed by using the mixed-fidelity design method with computational fluid dynamics equivalent-area analysis. The new inverse design process uses the reverse propagation of the pressure distribution from a mid-field location to a near-field location, converts the near-field into an equivalent-area distribution, generates a low-boom target for the reversed equivalent area of the configuration, and modifies the configuration to minimize the differences between the configuration’s reversed equivalent area and the low-boom target. The new inverse design process is used to modify a supersonic demonstrator concept for a cruise Mach number of 1.6 and a cruise weight of 30,000 lb. The modified configuration has a fully shaped ground signature that has a perceived loudness value of 78.5, whereas the original configuration has a partially shaped aft signature with a perceived loudness of 82.3.

Parametric study of the class-shape-refinement-transformation method

The application of a novel parametrization technique to the optimization of aircraft shapes is presented. This class-shape-refinement transformation (CSRT) technique combines an analytical function (class function), a set of Bernstein polynomials (shape function) and a B-spline (refinement function) and can be used to model various aircraft components. It allows for both global and local control of a shape and forms a very efficient and intuitive way of mathematically describing aircraft parts. A parametric study was performed that shows the behaviour of the shape as a function of a number of different parameters, such as total number of shape variables and Bernstein/B-spline coefficient ratio. The CSRT method was used to approximate a typical aircraft wing cross-section and the results showed a very non-linear relationship between the number of shape variables and the error of the approximation, expressed in terms of a correlation factor. This behaviour has been thoroughly analysed. Additionally, optimization results were obtained that show that the CSRT method was successfully coupled to an aerodynamic flow solver. The objective of the optimization runs was to maximize lift-to-drag ratio, but in principle any objective function could be used as long as its input follows from the aerodynamic analysis. The optimization algorithm is capable of largely removing the shock wave on an airfoil at a typical cruise Mach number.

Wavefront measurements of a laser-induced breakdown spark in still air

Experimental measurements of the wavefronts of the light from a laser-induced breakdown (LIB) spark in non-moving air are presented and compared to spark dimensional data acquired from photographic measurements of the spark. The data show that the variation in the spark emitted wavefront between ignitions can be directly related to the motion of the spark volumetric centroid. The dominant modal components of the emitted wavefront variations are presented, as well as quantitative results for the magnitude of the wavefront variations. The results are relevant to the use of LIB as a light source for the measurement of optical aberrations such as those caused by compressible (i.e., "aero-optic") flows around an aircraft in flight, and data are shown indicating that LIB could be successfully used to measure the aberrating effect of compressible shear layers and boundary layers at typical cruise Mach numbers.

Design Optimization of a Truss-Braced-Wing Transonic Transport Aircraft

the conventional cantilever configuration. One comparison produces a reduction of 45% in the fuel consumption while decreasing the minimum takeoff gross weight by 15%. For a second comparison, the fuel weight is reduced by 33% with a decreased minimum takeoff gross weight of 19%. Very attractive vehicle performance can be achieved without the necessity of decreasing cruise Mach number. The results also indicate that a truss-braced wing has a greater potential for improved aerodynamic performance than other innovative aircraft configurations. Further studieswillconsidertheinclusionofmorecomplextrusstopologiesandotherinnovativetechnologiesthatarejudged to be synergistic with truss-braced-wing configurations.

Inlet Stagnation Pressure Loss Estimate for a Simplified Thrust Model

In a previous work, the author presented a simplified thrust model to be used in predicting the performance of high-speed aircraft. That model contained an ad hoc estimate of the stagnation pressure loss in the inlet/diffuser of the engine. The current analysis attempts to capture some of the actual physics of the flow, modeling the incoming flow as a single oblique shock followed by a single normal shock, and including frictional losses in the subsonic portion of the diffuser. The result is a polynomial function of Mach number that captures much of the actual physics of the flow. Results using the new model are compared to the previous model and the data on which that model was based. A sample calculation of cruise Mach number for a jet aircraft is also included.

Controlled canard configuration study for a solid rocket motor based unmanned air vehicle

This paper presents an aerodynamic study for a target unmanned air vehicle with controlled canard, at cruise Mach number between 0.7 and 0.85, using a solid rocket motor as power. The results of semi-empirical method (DATCOM), computational fluid dynamics (CFD) were compared with wind tunnel tests to evaluate perdition accuracy. Here DATCOM shows big discrepancy in pitching-moment coefficient; the CFD estimation of static aerodynamic force and moment agree well with wind tunnel data. An analysis was focused on the longitudinal data for complicated controlled canard-elements interactions and canard vortex behavior. The canard vortex interaction with tail wing of initial configuration concept with a thick canard (NACA0012) and the final improved configuration with a thin canard (NACA0007) were reviewed. It is proved that the stability and control characteristics of whole TUAV are sensitive to the thickness of canard in transonic peed regime because of different canard vortex intensity and effect generated.

MHD Flow-Control for Hypersonic Flight

At the beginning of the seventies, experimental research using short duration supersonic shock driven wind tunnel were conducted and showed the e‐ciency of Lorentz force action on supersonic (Mach 1.4) dense (1 bar) low magnetic Reynolds number hot (10000 K) argon ∞ows. When working as a generator, the linear Faraday MHD converter was e‐cient enough, due to high electrical conductivity (3000 S/m) to create a front shock wave at the intake of the constant cross-section MHD converter. Accelerations up to 5000 m/s were obtained in 10 cm long channels. This suggested the possibility to achieve complete shock wave and turbulent wake cancellation, through MHD bypass concept. Such work gave publications in peer-reviewed journals and presentation in international MHD meetings (French Academy of Sciences 1970, Moscow 1983, Tsukuba 1986, Beijing 1992). The work was conducted both through computational simulations (based on the method of characteristics) and hydraulic simulation experiments. We present the synthesis of such studies through a project of a hypersonic vehicle based on an MHD bypass concept, landing under its own steam, using classical turbojets. Then, at Mach 3 and high altitude, MHD controlled inlets are opened. A wall converter slows down the hypersonic incoming air ∞ow, without excessive heating, feeding a ramjet system. The subsequent electric power provides an additional impulse to the ∞ow in the exhaust section. Cruise Mach number: 12. Local Hall parameter regime produces high voltage that, sent to the leading edges, creates a plasma cushion which prevents too high thermal ∞ux. Lift is provided by wave riding technique. Additional rocket propeller could transform it as a reusable space launcher.

Multipoint Optimization of a Short-Range Quiet Passenger Aircraft

In this paper, the wing shape of a mid-sized passenger aircraft for short-distance use is designed via a multi-point optimization at cruise and climb conditions. The cruise Mach number is set to 0.70: this enables to adopt a low-sweepback wing, which presents many advantages in terms of fuel efficiency, noise generation and manufacturing. The objective of the present work is to investigate the feasibility of the concept through an optimization approach. An Euler solver is used to analyze the aerodynamic performance of the wing. A Genetic Algorithm (GA) is used to search for optimal solutions over a Kriging approximation model. Two optimizations of the wing are conducted by changing the number of control sections to obtain high L/D at cruise and climb conditions. Finally, the optimal wing is attached to a fuselage and the model is analyzed. The Prandtl-Hoerner’s formula is used to assess the friction drag of the configuration. As a result, a reasonably high value of L/D is achieved. The results indicate that the concept of a low sweepback wing can represent an attractive option for many short-range flight routes.

巡航マッハ数に応じた超音速ジェットエンジンの多目的最適化

The aim of this paper is to apply a multi-objective optimization generic algorithm (MOGA) to the conceptual design of the hypersonic/supersonic vehicles with different cruise Mach number. The pre-cooled turbojet engine is employed as a propulsion system and some engine parameters such as the precooler size, compressor size, compression ratio and fuel type are varied in the analysis. The result shows that the optimum cruise Mach number is about 4 if hydrogen fuel is used. Methane fuel instead of hydrogen reduces the vehicle gross weight by 33% in case of the Mach 2 vehicle.