Journal of Aircraft

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ISSN / EISSN : 0021-8669 / 1533-3868
Total articles ≅ 10,590
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, Pallavi Rastogi, Ashish Bhatt
Journal of Aircraft pp 1-11;

Aircraft infrared (IR) signature studies are complex, due to their dependence on several parameters, e.g., line of sight (LOS) and viewing aspect (LOS^). There is no equivalent counterpart of the well-known radar cross section (RCS) that can be easily used for obtaining IR lock-on range, RLO,λ1−λ2. This study introduces the concept of IR cross section (IRCS)λ1−λ2 of complete aircraft as seen by seeker in λ1–λ2 band in LOS^ and of aircraft part in an orthogonal view. The IR solid angle (ωIR,λ1−λ2) subtended by (IRCS)λ1−λ2 of complete aircraft and of an aircraft part is also introduced, which is the basis for redefined general criterion for lock-on by IR-guided missile. Equality based on ωIR,λ1−λ2 at aircraft part level and at complete aircraft level is the basis for obtaining RLO,λ1−λ2 of complete aircraft in terms of its (IRCS)λ1−λ2. Dimensionless scalar factor, ΠIRCS,λ1−λ2, converting visual area to (IRCS)λ1−λ2 is studied, for temperatures ranging from close to ambient to maximum afterburning (reheat) mode of aeroengine. The ΠIRCS,λ1−λ2 enables assessment of effectiveness of IR seekers with single or dual bands, for all-aspect engagement and for rear-view only.
Jeffrey D. Taylor, Douglas F. Hunsaker
Journal of Aircraft pp 1-14;

As contemporary aerostructural research for aircraft design trends toward high-fidelity computational methods, aerostructural solutions based on theory are often neglected or forgotten. In fact, in many modern aerostructural wing optimization studies, the elliptic lift distribution is used as a reference in place of theoretical aerostructural solutions with more appropriate constraints. In this paper, the authors review several theoretical aerostructural solutions that could be used as reference cases for wing design studies, and these are compared to high-fidelity solutions with similar constraints. Solutions are presented for studies with 1) constraints related to the wing integrated bending moment, 2) constraints related to the wing root bending moment, and 3) structural constraints combined with operational constraints related to either wing stall or wing loading. It is shown that, under appropriate design constraints, theoretical solutions for the optimum lift distribution may capture aerostructural coupling sufficiently to serve as appropriate reference cases for higher-fidelity solvers. A comparison of theoretical and high-fidelity solutions for the optimum wingspan and corresponding drag reveals important insights into the effects of certain aerodynamic and structural parameters and constraints on the aerodynamic and structural coupling involved in aerostructural wing design and optimization.
Forrest L. Carpenter, Jackson Kassing, Paul G. A. Cizmas
Journal of Aircraft pp 1-17;

This paper presents the results of turbulent flow simulations prepared for the Third AIAA Sonic Boom Prediction Workshop. Solutions were prepared for two geometries that exhibited shock–plume interaction features: the NASA biconvex shock–plume interaction wind-tunnel model, and the NASA C608 low-boom concept aircraft. The unstructured, finite volume Navier–Stokes solver UNS3D was used to predict the turbulent near-field flow with Menter’s κ–ω shear stress transport turbulence model. Four solver setups with different combinations of the flux function, gradient reconstruction, and slope limiter were applied to each geometry. UNS3D predictions were compared against a grid-specific ensemble dataset created using workshop participant submissions and experimental data when available. Predictions of the biconvex geometry were found to be in good agreement with both the experimental and ensemble datasets. The weighted least-squares with QR decomposition (LSQR) gradient reconstruction approach was only successful for the biconvex geometry, with no converged cases being achieved on C608 grids. The Dervieux limiter, used only to aid simulations on the C608 using weighted LSQR, was observed to be too dissipative to be of practical use. Near-field predictions of the C608 geometry using Green–Gauss gradient reconstruction and a modified Venkatakrishnan limiter were found to correlate well with the ensemble data. The perceived level of the resulting sonic boom carpet was 1.5 dB quieter along the first 20 deg of the carpet than the ensemble mean, and it peaked at an azimuth angle of 30 deg with a value of 76.1 dB. A study of the undertrack near-field pressure signature revealed the components of the predicted signatures responsible for the variances observed in the sonic boom carpet.
Wade M. Spurlock, Michael J. Aftosmis, Marian Nemec
Journal of Aircraft pp 1-17;

Simulation results are presented for all cases from the Third AIAA Sonic Boom Prediction Workshop. An inviscid, embedded-boundary Cartesian-mesh flow solver is used in conjunction with adjoint-based mesh adaptation to compute near-field pressure signatures. Specialized techniques are applied to maximize accuracy and minimize cost on Cartesian meshes. Regions of the flow most sensitive to discretization error are identified using Richardson extrapolation. Timing results and mesh sizes for near-field cases demonstrate that decomposition into multiple off-track simulations is efficient in both computational time and wall-clock time, with results among the least computationally expensive of those presented at the workshop. Pressure signals are propagated to the ground using an augmented Burgers equation solver to predict boom carpets. Ground signatures and loudness metrics are presented for a standard atmosphere as well as an atmosphere with wind profiles that affect overall noise levels and can significantly widen the boom carpet. Mesh convergence studies show that high sampling frequencies, around 500 kHz, are required for on-track propagation; the sampling frequency increases at large off-track angles due to longer acoustic ray paths. Overall, the approach presented here yields accurate results for predicting low-sonic-boom signatures.
Timothy Chau, David W. Zingg
Journal of Aircraft pp 1-19;

The aerodynamic design and fuel burn performance of a Mach-0.78 strut-braced-wing regional jet is investigated through aerodynamic shape optimization based on the Reynolds-averaged Navier–Stokes equations. Conceptual-level multidisciplinary design optimization is first performed to size the strut-braced-wing aircraft for a design mission similar to the Embraer E190-E2, with a design range of 3100 nmi at a maximum capacity of 104 passengers, and a maximum payload of 30,200 lb. For direct performance comparisons, a conventional tube-and-wing regional jet is also sized and optimized based on the same reference aircraft. Gradient-based aerodynamic shape optimization is then performed on wing–body–tail models of each aircraft, with the objective of drag minimization at cruise over a 500 nmi nominal mission. Design variables include twist and section shape degrees of freedom, which are realized through a free-form and axial deformation geometry control system, whereas nonlinear constraints include constant lift, zero pitching moment, minimum wing volume, and minimum maximum thickness-to-chord ratios. Results indicate that the optimizer is capable of mitigating shock formation, boundary-layer separation, and other flow interference effects from each wing design, including those within the wing–strut junction of the strut-braced wing. With year 2020 technology levels, the strut-braced-wing regional jet offers a 12.9% improvement in cruise lift-to-drag ratio over an Embraer E190-E2-like conventional tube-and-wing aircraft, which translates to a 7.6% reduction in block fuel for the nominal mission.
James L. Lankford, Inderjit Chopra
Journal of Aircraft pp 1-19;

Instantaneous force and structural deformation experiments were performed on a flexible, structurally characterized, low-aspect-ratio representative flapping wing. A six-component force balance was used to measure aerodynamic-force time history over a flap cycle. A VICON motion capture setup tracked the wing deformations while flapping. Measured aerodynamic forces and wing deformations were compared against coupled computational fluid dynamics/computational structural dynamics (CFD/CSD) aeroelastic analysis results. The CFD solver is an unsteady Reynolds-averaged Navier–Stokes solver. The CSD solver is a general-purpose multibody dynamics solver capable of modeling geometrically nonlinear beam and shell elements. The CFD/CSD results were able to capture the overall trend in aerodynamic forces and wing deformations. Although predicted and measured variations in lift were similar, the drag-force magnitudes tended to be underpredicted by the coupled aeroelastic solver. The coupled CFD/CSD solver was used to evaluate the influence of wing flexibility on wing performance. Results showed that, in particular instances, decreasing wing stiffness increased the time-averaged aerodynamic lift and lift-to-power ratio. Decreased wing stiffness led to reduced leading-edge vortex circulation strength. This suggests that the increased wing performance is mainly due to a redirection of the resultant force vector allowed by the increased wing compliance.
James R. Gibson, Virginia Mitchell
Journal of Aircraft pp 1-10;

Tiltrotor aircraft ship suitability and envelope expansion testing aboard a U.S. Navy aircraft carrier is discussed. The purpose was to validate and expand day/night vertical launch and recovery wind envelopes, and the development of short takeoff and minimum run-on landing wind envelopes. The preparation phase incorporated a six-degree-of-freedom motion flight simulator to evaluate flight dynamics prior to on-shore testing using V-22 developmental and production aircraft. The short takeoff and minimum run-on landing testing totaled 12.3 flight hours with 33 takeoffs and 32 landings. The gross weights ranged from 51,000 to 57,000 lb with ±15 deg of relative wind from 0 to 45 kt. The test pilots provided assessments of handling and flying qualities throughout the maneuvers using the deck interface pilot effort scale. Predictability tests were required to determine the pilot’s ability to land within the defined touchdown zone and speeds for safely stopping within the braking zone. The results show the integrated nature of shipboard flight testing and the process of combining pilot workload ratings, control strategies, and aerodynamic analyses to field a needed capability.
Hiroaki Ishikawa, Shinya Koganezawa, Yoshikazu Makino
Journal of Aircraft pp 1-10;

The Third AIAA Sonic Boom Prediction Workshop was held in January 2020, and in the near-field section of the workshop, several simulated near-field pressure signatures of the C608 airplane, an early version of the NASA X-59, were compared. The authors have applied the unstructured/structured overset grid method to the C608 airplane case. In this method, an unstructured grid is used in the vicinity of the airplane, and the flowfield at several body lengths away from the airplane is simulated by oversetting a structured Mach cone aligned grid with the unstructured grid. It is found that the near-field pressure signatures can be obtained accurately with sharp pressure jumps in the signature by using a fine structured grid in the midfield even when a relatively coarse unstructured grid is used in the near-field of the airplane. The results show that the unstructured/structured overset grid method is a way to obtain accurate near-field pressure signatures for sonic boom estimation efficiently.
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