Two variants of a conceptual, Mach 1.5 low-boom airframe are developed using a nonlinear, inviscid flow solver coupled with an adjoint-based shape optimization framework. In the first variant, a wing planform with a large trailing edge extension shields the undertrack observer from the nozzle shock system. The size of this extension is reduced on the second variant, allowing unobstructed passage of the nozzle shock. Each variant is evaluated in both powered and flow-through states using high-resolution meshes constructed using an adaptive, adjoint-driven approach. Conventional converging-diverging nozzles are examined at cruise and full-reheat thrust, while two plug nozzles with di↵ering cone angles are examined at the cruise thrust condition. For each case, performance is characterized in terms of the e↵ect on the near-field pressure signal, the propagated ground signature, and the predicted loudness level at the ground. The results show that for this configuration, shielding e↵ectively controls the sonic boom level at the ground regardless of nozzle power state. In the absence of shielding, the reduced flow-field disturbance due to a plug nozzle is shown to o↵er only slightly reduced low-boom performance relative to the un-powered baseline. NomenclatureJ Shape optimization objective J r Mesh refinement objective A Cross-sectional area, ft 2 A e Equivalent area distribution, ft 2 b Superellipse segment height, ft C D Drag coe cient C L Lift coe cient C p Pressure coe cient h Vehicle nose height above near-field sensor, ft L E↵ective length, ft L/D Lift-to-drag ratio L s Nozzle plug support sting length, ft M Mach number n Superellipse exponent p Static pressure r Radius, ft r f Nozzle plug fillet radius, ft r p Nozzle plug radius, ft r s Nozzle plug support sting radius, ft S Near-field sensor length, ft s Distance along near-field sensor, ft S ref Reference area, ft 2

The desire to reduce or eliminate the operational restrictions of supersonic aircraft over populated areas has led to extensive research at NASA. Restrictions are due to the disturbance of the sonic boom, caused by the coalescence of shock waves formed by the aircraft. A study has been performed focused on reducing the magnitude of the sonic boom N-wave generated by airplane components with a focus on shock waves caused by the exhaust nozzle plume. Testing was completed in the 1-foot by 1-foot supersonic wind tunnel to study the effects of an exhaust nozzle plume and shock wave interaction. The plume and shock interaction study was developed to collect data for computational fluid dynamics (CFD) validation of a nozzle plume passing through the shock generated from the wing or tail of a supersonic vehicle. The wing or tail was simulated with a wedgeshaped shock generator. This test entry was the first of two phases to collect schlieren images and off-body static pressure profiles. Three wedge configurations were tested consisting of strut-mounted wedges of 2.5degrees and 5-degrees. Three propulsion configurations were tested simulating the propulsion pod and aft deck from a low boom vehicle concept, which also provided a trailing edge shock and plume interaction. Findings include how the interaction of the jet plume caused a thickening of the shock generated by the wedge (or aft deck) and demonstrate how the shock location moved with increasing nozzle pressure ratio. Nomenclature D= test nozzle outer diameter, inches NPR = nozzle pressure ratio, Pt / P∞ P = local static pressure, psia Pt = total pressure in nozzle, psia P∞ = free stream static pressure, psia ΔP/P = normalized static pressure, (P -P∞)/ P∞ x = axial distance from jet simulator nozzle exit plane, inches y = vertical distance from nozzle centerline, inches

Computational fluid dynamics (CFD) analysis has been performed to study the plume effects on sonic boom signature for isolated nozzle configurations. The objectives of these analyses were to provide comparison to past work using modern CFD analysis tools, to investigate the differences of high aspect ratio nozzles to circular (axisymmetric) nozzles, and to report the effects of underexpanded nozzle operation on boom signature. CFD analysis was used to address the plume effects on sonic boom signature from a baseline exhaust nozzle. Near-field pressure signatures were collected for nozzle pressure ratios (NPRs) between 6 and 10. A computer code was used to extrapolate these signatures to a ground-observed sonic boom N-wave. Trends show that there is a reduction in sonic boom N-wave signature as NPR is increased from 6 to 10. The performance curve for this supersonic nozzle is flat, so there is not a significant loss in thrust coefficient as the NPR is increased. As a result, this benefit could be realized without significant loss of performance. Analyses were also collected for a high aspect ratio nozzle based on the baseline design for comparison. Pressure signatures were collected for nozzle pressure ratios from 8 to 12. Signatures were nearly twice as strong for the two-dimensional case, and trends also show a reduction in sonic boom signature as NPR is increased from 8 to 12. As low boom designs are developed and improved, there will be a need for understanding the interaction between the aircraft boat tail shocks and the exhaust nozzle plume. These CFD analyses will provide a baseline study for future analysis efforts. NomenclatureC fg = computed thrust (from CFD)/ideal thrust NPR = nozzle pressure ratio = P t /P ∞ P = local static pressure, psia P t = total pressure in nozzle P ∞ = free-stream static pressure ΔP/P ∞ = (P -P ∞ )/P ∞ ΔP = P -P ∞ x = axial distance, in. D = test nozzle diameter, in. x/D = nondimensional axial distance from jet simulator nose cone

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In late 2015, NASA's Aeronautics Research Mission Directorate (ARMD) funded an experiment in rapid design and rapid teaming to explore new approaches to solving challenging design problems in aeronautics in an effort to cultivate and foster innovation. This report summarizes several lessons learned from the rapid design portion of the study. This effort entailed learning and applying design thinking, a human-centered design approach, to complete the conceptual design for an open-ended design challenge within six months.The design challenge focused on creating a capability to advance experimental testing of autonomous aeronautics systems, an area of great interest to NASA, the US government as a whole, and an entire ecosystem of users and developers around the globe. A team of nine civil servant researchers from three of NASA's aeronautics field centers with backgrounds in several disciplines was assembled and rapidly trained in design thinking under the guidance of the innovation and design firm IDEO. The design thinking process, while used extensively outside the aerospace industry, is less common and even counter to many practices within the aerospace industry. In this report, several contrasts between common aerospace research and development practices and design thinking are discussed, drawing upon the lessons learned from the NASA rapid design study.The lessons discussed included working towards a design solution without a set of detailed design requirements, which may not be practical or even feasible for management to ascertain for complex, challenging problems. This approach allowed for the possibility of redesigning the original problem statement to better meet the needs of the users. Another lesson learned was to approach problems holistically from the perspective of the needs of individuals that may be affected by advances in topic area instead of purely from a technological feasibility viewpoint. The interdisciplinary nature of the design team also provided valuable experience by allowing team members from different technological backgrounds to work side-by-side instead of dividing into smaller teams, as is frequently done in traditional multidisciplinary design. The team also learned how to work with qualitative data obtained primarily through the 70-plus interviews that were conducted over the course of this project, which was a sharp contrast to using quantitative data with regards to identifying, capturing, analyzing, storing, and recalling the data. When identifying potential interviewees who may have useful contributions to the design subject area, the team found great value in talking to non-traditional users and potential beneficiaries of autonomous aeronautics systems whose impact on the aeronautics autonomy ecosystem is growing swiftly. Finally, the team benefitted from using "sacrificial prototyping," which is a method of rapidly prototyping draft concepts and ideas with the intent of enabling potential users to provide significant feedback early in the design process. This contrasts the mo...