Aircraft flown in formation can realize significant reductions in drag by flying in regions of wake upwash. However, most transports fly at transonic speeds where the impact of compressibility on formation flight is not well understood. This study uses an Euler solver to analyze the inviscid aerodynamic forces and moments of transonic wing/body configurations flying in a two-aircraft formation. Formations with large streamwise separation distances (10–50 wingspans) are considered. This work indicates that compressibility-related drag penalties in formation flight may be eliminated by slowing 2–3% below the nominal out-of-formation cruise Mach number, either at fixed lift coefficient or fixed altitude. The latter option has the additional benefit that the aerodynamic performance of the formation improves slightly at higher lift coefficients. Although optimal in-formation lift coefficients are not as high as those estimated by incompressible analyses, modest increases in altitude can yield further improveme...

Stage separation is a critical technical issue for developing two-stage-to-orbit (TSTO) launch systems with widebody carrier aircraft that use air-breathing propulsion and launch vehicle stages that use rocket propulsion. During conceptual design phases, this issue can be addressed with a combination of engineering methods, computational fluid dynamics simulations, and trajectory analysis of the mated system and the launch vehicle after staging. The outcome of such analyses helps to establish the credibility of the proposed TSTO system and formulate a ground-based test programme for the preliminary design phase. This approach is demonstrated with an assessment of stage separation from the shuttle carrier aircraft. Flight conditions are determined for safe mated flight, safe stage separation, and for the launch vehicle as it commences ascending flight. Accurate assessment of aerodynamic forces and moments is critical during staging to account for interference effects from the proximities of the two large vehicles. Interference aerodynamics have a modest impact on the separation conditions and separated flight trajectories, but have a significant impact on the interaction forces.

This paper presents the findings of a study conducted tn 2010 by the NASA Innovation Fund Award project entitled "Elastically Shaped Future Air Vehicle Concept". The study presents three themes in support of meeting national and global aviation challenges of reducing fuel burn for present and future aviation systems. The first theme addresses the drag reduction goal through innovative vehicle configurations via non-planar wing optimization. Two wing candidate concepts have been identified from the wing optimization: a drooped wing shape and an inflected wing shape. The drooped wing shape is a truly biologically inspired wing concept that mimics a seagull wing and could achieve about 5% to 6% drag reduction, which is aerodynamically significant. From a practical perspective, this concept would require new radical changes to the current aircraft development capabilities for new vehicles with futuristic-looking wings such as this concept. The inflected wing concepts could achieve between 3% to 4% drag reduction. While the drag reduction benefit may be less, the inflected-wing concept could have a near-term impact since this concept could be developed within the current aircraft development capabilities. The second theme addresses the drag reduction goal through a new concept of elastic wing shaping control. By aeroelastically tailoring the wing shape with active control to maintain optimal aerodynamics, a significant drag reduction benefit could be realized. A significant reduction in fuel burn for long-range cruise from elastic wing shaping control could be realized. To realize the potential of the elastic wing shaping control concept, the third theme emerges that addresses the drag reduction goal through a new aerodynamic control effector called a variable camber continuous trailing edge flap. Conventional aerodynamic control surfaces are discrete independent surfaces that cause geometric discontinuities at the trailing edge region. These discontinuities promote vorticities which result in drag rises as well as noise sources. The variable camber trailing edge flap concept could provide a substantial drag reduction benefit over a conventional discrete flap system. Aerodynamic simulations show a drag reduction of over 50% could be achieved with the flap concept over a conventional discrete flap system.

Flying airplanes in extended formations, with separation distances of tens of wingspans, significantly improves safety while maintaining most of the fuel savings achieved in close formations. The present study investigates the impact of roll trim and compressibility at a fixed lift coefficient on the benefits of extended-formation flight. An Euler solver with adjoint-based mesh refinement combined with a wake propagation model is used to analyze a two-body echelon formation at a separation distance of 30 spans. Two geometries are examined: a simple wing and a wing-body geometry. Energy savings, quantified by both formation drag fraction and span efficiency factor, are investigated at subsonic and transonic speeds for a matrix of vortex locations. The results show that, at fixed lift and trimmed for roll, the optimal location of vortex impingement is about 10% inboard of the trailing airplane's wing tip. Interestingly, the improvement in drag fraction is relatively robust in the vicinity of the optimal position. Over 90% of energy benefits can be obtained with a 5% variation in vertical and 10% variation in spanwise positions. Control surface deflections required to achieve roll trim reduce the benefits of formation flight by 3-5% at subsonic speeds and 9-11% at transonic speeds. Overall, simulations show peak induced drag saving for the trail aircraft are 54% in subsonic flow and 35% in transonic flow while accounting for trim.

A procedure for validating the capability of Over ow for predicting the plume e ects on the Ares I launch vehicle during the stage separation process is reported. There are no wind tunnel or ight data available for the Ares I con gurations with the full set of plumes active during the stage separation scenario. Instead a series of \representative unit problems" will be used for validating Over ow. Four di erent methods are used in this paper. The numerical predictions from Over ow are compared to exact solutions, experimental data, ight data from the Apollo era, and nally numerical predictions from other Navier-Stokes codes. The numerical predictions of plumes e ects are in uenced by grid resolution, turbulence models, plume gas properties, ow conditions, boundary conditions, geometric parameters, and ow unsteadiness. All of these have been investigated and the results are summarized here.

This paper conducts a numerical study of four sets of aerodynamic control surfaces, called grid fins, mounted on a full Launch Abort Vehicle geometry using a Cartesian Euler Solver with embedded boundaries and adjoint-driven adaptive meshing. Since Cartesian methods are insensitive to complex geometry, this makes them suitable for the highly complex resulting configuration. Numerical results are compared against subtransand supersonic wind tunnel data in order to examine our ability to accurately predict the force and moment increments afforded by these unconventional control surfaces. A database of 1152 separate cases were ran including 12 different Mach numbers from 0.5 2.5, 15 different angle of attacks (0 15◦), and 6 different geometries. Overall, the simulation data show good agreement with tunnel runs and similar rankings of configurations and trends are found between wind tunnel and simulation results suggesting a large potential for Cartesian Euler Solvers in accurately predicting force and moment increments for grid fins on the Launch Abort Vehicle. In particular, pitching moment was predicted accurately over the entire Mach-alpha space.

The extent of boundary-layer separation due to the presence of a jet plume was studied for a cone-cylinder-flare configuration using the Overset Navier-Stokes CFD code, OVERFLOW. Simulations were made at a free-stream Mach number of 4.65 for nozzle exit pressure to free-stream pressure ratios ranging from jet off to 168. CFD results were obtained assuming a fully laminar boundary layer, fully turbulent boundary layer, and a laminar-turbulent transitional boundary layer. The one-equation Spalart-Allmaras (SA) and the two-equation SST model were compared for the turbulent cases. The effect of wall boundary conditions was studied by comparing the isothermal wall to the adiabatic wall conditions. CFD results were compared to existing experimental data obtained from the Langley Unitary Plan wind tunnel. The results indicate strong plume-induced flow separation when modeled as fully laminar and weak separation, if at all, when modeled as fully turbulent. The best agreement with experimental results was achieved when the trip location was placed near the end of the flare. The adiabatic wall boundary condition best suited the wind tunnel experiment. However predicting the extent of the plume induced flow separation with numerical simulations is still unreliable due to the extreme sensitivity of the flow field to turbulence models, boundary conditions, and nozzle pressure ratios.

Aerodynamic characterization in the ascent phase is one of the necessary steps to designing a successful rocket. Computational fluid dynamics (CFD) simulations that employ the structured overset grid philosophy are commonly used for aerodynamics analysis. While the method is highly desirable due to its ability to provide viscous flow solutions for complex geometries, the preliminary geometry processing work required to generate grids is a major bottleneck to efficiently obtaining fluid simulation results. The present paper proposes strategies to improve the grid generation process by eliminating the mundane tasks that can be automated with limited knowledge of the geometry and little user input. The automation is targeted for rocket bodies and the protuberances that are commonly placed on rockets. However, the resulting tools may be applicable to other geometries. The present paper’s focus is on a scripting framework to automate surface and volume mesh generation from a native computer aided design (CAD) solid model geometry definition. The resulting process is able to generate overset meshes with fewer lines of code and less user input culminating in savings in time to process a clean solid model CAD geometry to a CFD-ready mesh.