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NASA NTRS1993

Techniques for estimating Space Station aerodynamic characteristics

A method was devised and calculations were performed to determine the effects of reflected molecules on the aerodynamic force and moment coefficients for a body in free molecule flow. A procedure was developed for determining the velocity and temperature distributions of molecules reflected from a surface of arbitrary momentum and energy accommodation. A system of equations, based on momentum and energy balances for the surface, incident, and reflected molecules, was solved by a numerical optimization technique. The minimization of a 'cost' function, developed from the set of equations, resulted in the determination of the defining properties of the flow reflected from the arbitrary surface. The properties used to define both the incident and reflected flows were: average temperature of the molecules in the flow, angle of the flow with respect to a vector normal to the surface, and the molecular speed ratio. The properties of the reflected flow were used to calculate the contribution of multiply reflected molecules to the force and moments on a test body in the flow. The test configuration consisted of two flat plates joined along one edge at a right angle to each other. When force and moment coefficients of this 90 deg concave wedge were compared to results that did not include multiple reflections, it was found that multiple reflections could nearly double lift and drag coefficients, with nearly a 50 percent increase in pitching moment for cases with specular or nearly specular accommodation. The cases of diffuse or nearly diffuse accommodation often had minor reductions in axial and normal forces when multiple reflections were included. There were several cases of intermediate accommodation where the addition of multiple reflection effects more than tripled the lift coefficient over the convex technique.

Thomas, Richard E.↗

NASA NTRS1996

Evolution of the revolutionary blended-wing-body

The Blended-Wing-Body (BWB) airplane concept represents a potential revolution in subsonic transport efficiency for Very Large Airplanes (VLA's). NASA is sponsoring an advanced concept study to demonstrate feasibility and begin development of this new class of airplane. In this study, 800 passenger BWB and conventional configuration airplanes have been compared for a 7000 nautical mile design range, where both airplanes are based on technology keyed to 2015 entry into service. The BWB has been found to be superior to the conventional configuration in the following areas: Fuel Burn--31% lower, Takeoff Weight -- 1 3% lower, Operating Empty Weight -- 10% lower, Total Thrust -- 16% lower, and Lift/Drag --35% higher. The BWB advantage results from a double deck cabin that extends spanwise providing structural and aerodynamic overlap with the wing. This reduces the total wetted area of the airplane and allows a high aspect ratio to be achieved, since the deep and stiff centerbody provides efficient structural wingspan. Further synergy is realized through buried engines that ingest the wing's boundary layer, and thus reduce effective ram drag. Relaxed static stability allows optimal span loading, and an outboard leading-edge slat is the only high-lift system required.

Liebeck, Robert H.↗

NASA NTRS1997

Future Aerospace Technology in the Service of the Alliance: Sustained Hypersonic Flight - Volume 3

Sustained Hypersonic Flight capability is envisioned for several future military and civilian applications, such as long range immediate reaction reconnaissance, high speed interception of air targets, long range precision strike against hardened or time critical targets, and access to space. The inherent reduction in time-to-target and low vulnerability will permit new operational tactics. The symposium outlined mission needs and operational scenarios for hypersonic vehicles. Addressed were: Technological issues and challenges in external hypersonic aerodynamics and design, propulsion and engine/airframe integration, military ramjet applications, overall system design including structures and materials development needs, and test facilities. Ground based test facilities cannot provide full simulation at Mach numbers above 5. Therefore, the use of Computational Fluid Dynamics design tools is essential. For routine use of the computational tools in analysis, design and optimization, it is necessary to reduce the time for the entire computational process by two orders of magnitude. Hypersonic boundary layer transition remains a critical design issue because of the important impact on engine drag and on heating, which can affect the choice of materials and thermal protection systems. In addition, reliable prediction of scramjet net thrust is an absolute must in resolving hypersonic air breathing vehicle design issues. Due to current facility and computational shortfalls, the development of future hypersonic flight systems requires research flight tests in the technology areas of boundary layer transition and air-breathing propulsion engine performance. For sustained hypersonic flight beyond Mach 6, the supersonic combustion ramjet (scramjet) engine is the only choice for the near future. Only this air-breathing concept offers a significant promise of large reductions in required propellant fractions, increased payload fractions, and reduced size vehicles, together with a foreseeable technological feasibility. Airframe/engine integration, combustor design and thermal management are the predominant engineering tasks. Fuels, hydrogen or hydrocarbon, must be matched to the operational needs of military or civil use. Experience in existing ramjet propelled missiles capable of speeds up to Mach 4 can support the development effort. The potential mission and cost benefits of sustained hypersonic flight to both military and civil applications are tremendous. From the budget point of view, the possibility of sharing development costs between military and civil programs offers a specific advantage.

Source record↗

NASA NTRS2018

Tailored Wingbox Structures through Additive Manufacturing: A Summary of Ongoing Research at NASA LaRC

The use of wingbox structural design for improved performance (i.e., fuel burn reduction) of subsonic transports is driven by two trends: reduced structural weight and increased wingspan. These two trends are in direct competition, as the increased span will exacerbate the structural reaction to aerodynamic loading, and the reduced structural weight will nominally weaken the aircraft’s ability to handle this response. Novel structural configurations, enabled by recent improvements in manufacturing, may be critical toward bridging this gap.This paper summarizes pertinent activities at the NASA Langley Research Center in terms of additive manufacturing of metallic wing structures and substructures. Numerical design optimization activities are summarized as well, in order to understand where on a wingbox an additively-manufactured part may be useful and the way in which that part beneficially impacts the flight physics. The paper concludes with a discussion of how these two research paths may be better married in order to fully integrate both the benefits and realistic limitations of additive manufacturing and numerical structural design.

Stanford, Bret. K.↗

NASA NTRS1986

Two Programs for Supersonic Wing Design and Analysis

COREL and W12SC3 useful in aerodynamic design and analysis of wings for supersonic speeds. COREL (Conical Relaxation) program solves nonlinear full potential equation for spanwise section of wing in crossflow plane, and option exists to correct result for nonconical geometry. W12SC3 applies lineartheory panel methods to compute solutions for wing/body configuration. Programs restricted to supersonic flows and useful for many design, analysis, and optimization applications. COREL and W12SC3 written in FORTRAN IV for batch execution.

Mason, W. H.↗

NASA NTRS1999

Integrated Computational System for Aerodynamic Steering and Visualization

In February of 1994, an effort from the Fluid Dynamics and Information Sciences Divisions at NASA Ames Research Center with McDonnel Douglas Aerospace Company and Stanford University was initiated to develop, demonstrate, validate and disseminate automated software for numerical aerodynamic simulation. The goal of the initiative was to develop a tri-discipline approach encompassing CFD, Intelligent Systems, and Automated Flow Feature Recognition to improve the utility of CFD in the design cycle. This approach would then be represented through an intelligent computational system which could accept an engineer's definition of a problem and construct an optimal and reliable CFD solution. Stanford University's role focused on developing technologies that advance visualization capabilities for analysis of CFD data, extract specific flow features useful for the design process, and compare CFD data with experimental data. During the years 1995-1997, Stanford University focused on developing techniques in the area of tensor visualization and flow feature extraction. Software libraries were created enabling feature extraction and exploration of tensor fields. As a proof of concept, a prototype system called the Integrated Computational System (ICS) was developed to demonstrate CFD design cycle. The current research effort focuses on finding a quantitative comparison of general vector fields based on topological features. Since the method relies on topological information, grid matching and vector alignment is not needed in the comparison. This is often a problem with many data comparison techniques. In addition, since only topology based information is stored and compared for each field, there is a significant compression of information that enables large databases to be quickly searched. This report will (1) briefly review the technologies developed during 1995-1997 (2) describe current technologies in the area of comparison techniques, (4) describe the theory of our new method researched during the grant year (5) summarize a few of the results and finally (6) discuss work within the last 6 months that are direct extensions from the grant.

Hesselink, Lambertus↗

NASA NTRS1995

Sensitivity analysis in multipole-accelerated panel methods for potential flow

In the design of an airframe, the effect of changing the geometry on resulting computations is necessary for design optimization. The geometry is defined in terms of a series of design variables, including design variables to define the wing planform, tail, canard, pylon, and nacelle. Design optimization in this research is based on how these design variable affect the potential flow. The potential flow is computed as a function of the geometry and location of a series of panels describing the airframe, which are in turn a function of the design variables. Multipole accelerated panel methods improve the computational complexity of the problem and thus are an attractive approach. To utilize the methods in design optimization, it was necessary to define the appropriate sensitivity derivatives. The overhead incurred from finding the sensitivity derivatives in conjunction with the original computation should be small. This research developed the background for multipole-accelerated panel methods and the framework for finding sensitivity derivatives in the methods. Potential flow panel codes are commonly used for powered-lift aerodynamic predictions for three dimensional geometries. Given an airframe which has been discretized into a series of panels to define the airframe geometry, potential is computed as a function of the influence of all panels on all other panels. This is a computationally intensive problem for which efficient solutions are desired to improve the computational time and to allow greater resolution by use of more panels. One such solution is the use of hierarchical multipole methods which entail approximations of the effects of far-field terms. Hierarchical multipole methods have become prevalent in molecular dynamics and gravitational physics, and have been introduced into the fields of capacitance calculations, computational fluid dynamics, and electromagnetics. The methods utilize multipole expansions to describe the effect of bodies (i.e. particles, astrophysical bodies, panels, etc.) within a sphere on points distant from the sphere, where the influence diminishes as a function of distance. The expansions are exact with infinite series, however, for practical computations, the series are truncated and accuracy is selected based on the number of terms retained in the expansions. A hierarchical tree structure groups bodies together based on proximity to allow definition of multipole expansions for each group. The multipole expansions are then used to compute the effect of the bodies in a group on distant bodies.

Leathrum, James F., Jr.↗

NASA NTRS1991

Application of multidisciplinary optimization methods to the design of a supersonic transport

An optimization design method is discussed. This method is based on integrating existing disciplinary analysis and sensitivity analysis techniques by means of generalized sensitivity equations. A generic design system implementing this method is described. The system is being used to design the configuration and internal structure of a supersonic transport wing for optimum performance. This problem combines the disciplines of linear aerodynamics, structures, and performance. Initial results which include the disciplines of aerodynamics and structures in a conventional minimum weight design under static aeroelastic constraints are presented.

Barthelemy, J.-F. M.↗

NASA NTRS1991

Analysis and design of planar and non-planar wings for induced drag minimization

Improvements in the aerodynamic efficiency of commercial transport aircraft will reduce fuel usage with subsequent reduced cost, both monetary and environmental. To this end, the current research is aimed at reducing the overall drag of these aircraft with specific emphasis on reducing the drag generated by the lifting surfaces. The ultimate goal of this program is to create a wing design methodology which optimizes the geometry of the wing for lowest total drag within the constraints of a particular design specification. The components of drag which must be considered include profile drag, and wave drag. Profile drag is dependent upon, among other things, the airfoil section and the total wetted area. Induced drag, which is manifested as energy left in the wake by the trailing vortex system is mostly a function of wing span, but also depends on other geometric wing parameters. Wave drag of the wing, important in the transonic flight regime, is largely affected by the airfoil section, wing sweep, and so forth. The optimization problem is that of assessing the various parameters which contribute to the different components of wing drag, and determining the wing geometry which generates the best overall performance for a given aircraft mission. The primary thrust of the research effort to date was in the study of induced drag. Results from the study are presented.

Straussfogel, Dennis M.↗

NASA NTRS2000

Flexible Wing Model for Structural Sizing and Multidisciplinary Design Optimization of a Strut-Braced Wing

This paper describes a structural and aeroelastic model for wing sizing and weight calculation of a strut-braced wing. The wing weight is calculated using a newly developed structural weight analysis module considering the special nature of strut-braced wings. A specially developed aeroelastic model enables one to consider wing flexibility and spanload redistribution during in-flight maneuvers. The structural model uses a hexagonal wing-box featuring skin panels, stringers, and spar caps, whereas the aerodynamics part employs a linearized transonic vortex lattice method. Thus, the wing weight may be calculated from the rigid or flexible wing spanload. The calculations reveal the significant influence of the strut on the bending material weight of the wing. The use of a strut enables one to design a wing with thin airfoils without weight penalty. The strut also influences wing spanload and deformations. Weight savings are not only possible by calculation and iterative resizing of the wing structure according to the actual design loads. Moreover, as an advantage over the cantilever wing, employment of the strut twist moment for further load alleviation leads to increased savings in structural weight.

Gern, Frank H.↗

NASA NTRS1992

Application of multidisciplinary optimization methods to the design of a supersonic transport

A new optimization based design method is discussed. This method is based on integrating existing disciplinary analysis and sensitivity analysis techniques by means of generalized sensitivity equations. A generic design system implementing this method is described. The system is being used to design the configuration and internal structure of a supersonic transport wing for optimum performance. This problem combines the disciplines of linear aerodynamics, structures, and performance. Initial results which include the disciplines of aerodynamics and structures in a conventional minimum weight design under static aeroelastic constraints are presented.

Barthelemy, J.-F. M.↗

NASA NTRS2013

Parachute Models Used in the Mars Science Laboratory Entry, Descent, and Landing Simulation

An end-to-end simulation of the Mars Science Laboratory (MSL) entry, descent, and landing (EDL) sequence was created at the NASA Langley Research Center using the Program to Optimize Simulated Trajectories II (POST2). This simulation is capable of providing numerous MSL system and flight software responses, including Monte Carlo-derived statistics of these responses. The MSL POST2 simulation includes models of EDL system elements, including those related to the parachute system. Among these there are models for the parachute geometry, mass properties, deployment, inflation, opening force, area oscillations, aerodynamic coefficients, apparent mass, interaction with the main landing engines, and off-loading. These models were kept as simple as possible, considering the overall objectives of the simulation. The main purpose of this paper is to describe these parachute system models to the extent necessary to understand how they work and some of their limitations. A list of lessons learned during the development of the models and simulation is provided. Future improvements to the parachute system models are proposed.

Cruz, Juan R.↗

NASA NTRS2019

Airfoil Selection for Mars Rotor Applications

This short white paper is aimed at providing an overview of design considerations for airfoil choices for rotor applications in the Martian atmosphere, at very low chord-based Reynolds number flows, around Rec = O (103 -104). The low Reynolds number typical of rotorcraft operation in the Martian atmosphere reduces the rotor lifting force and efficiency, which is only partially compensated for by a lower gravity on Mars. Additionally, the low temperature and largely CO2 based atmosphere of Mars compound the overall aerodynamic problem by resulting in a lower speed of sound, further constraining rotor operation in the Martian atmosphere by limiting the maximum rotor tip speed possible so as not to exceed an acceptable tip Mach number. In light of the expected reduced rotor efficiency, evaluating airfoils for compressible, low-Reynolds number Mars rotor applications is key. Prior research on airfoil optimization and performance evaluation at low Reynolds numbers, especially in the compressible regime, is scarce and deserves further understanding. Specifically, the proposed goal stemming from this overview is to develop airfoils tailored to the unique demands of the second generation of Mars rotorcraft, i.e. the Mars Science Helicopter (MSH). This research focuses on the airfoil performance at low Reynolds numbers and hopes to add to the work performed by, amongst others, Kroo et al., Kunz and Kroo, Oyama and Fujii, Anyoji et al.

Koning, Witold J. F.↗

NASA NTRSDate not supplied

Advanced Parametric Lobe Mixer Concepts for Internally Mixed Nozzles

A continuation of efforts to optimize a dual-stream lobe mixer for supersonic commercial flight are detailed, adding scarfing and scalloping features to a baseline lobe mixer previously optimized within a baseline parametric space. These "advanced features" are parameterized by four quantities: scarfing angle 𝜁, scallop height ℎ 𝑠 , scallop depth 𝑑 𝑠 , and scallop conic focal parameter 𝜌 𝑠 . Two baseline mixers, one parallel and one vortical, form the basis for an eighteen-case study to investigate the effects of each parameter in single-axis fashion. Aerodynamic performance is evaluated through velocity coefficient 𝐶 𝑉 . Moderate performance increases are seen with increasing scallop height and depth on the parallel mixer, and with increasing scarfing angle on the vortical mixer. Flowfield analysis is presented through visualization of axial and cross-stream vorticity development through the duct, as well as density variation. A correlation between axial vorticity and density variation decay was found, with earlier introduction of axial vorticity found to result in improved aerodynamic performance and potential noise suppression.

Lobe Mixer↗

NASA NTRSDate not supplied

Aerocapture Trajectory Design for Uranus Orbiter

Introduction: The recently released National Academies Planetary Science and Astrobiology Decadal Survey 2023-2032 [1] identified the Ice Giants as the top priority science destination. While the survey acknowledged the potential for either a Uranus Orbiter and Probe (UOP) mission or a Neptune-Triton Odyssey mission, it ultimately identified the former as the highest priority new flagship mission. UOP missions calls for a launch window of opportunity between 2031-2038 with 12-15 year interplanetary cruise time along with a fully-propulsive Uranus Orbit Insertion burn on the order of a few km/s. However, a mission to Uranus with the same science payload could utilize aerocapture for orbit insertion to achieve both a significant reduction in the interplanetary cruise time and reduction in propulsive burn costs. Why Aerocapture: Aerocapture is a promising propellant and time-saving orbital insertion technique for planetary destinations with an atmosphere. Although not flight-proven, previous aerocapture systems studies in the literature have demonstrated both the validity and robustness of the technique at various planetary destinations. With respect to the Ice Giant planets, Neptune has seen more of the analysis in the literature. For science missions at Neptune, aerocapture can enable 1.4 times more delivered mass to orbit than an all-propulsive mission for the same launch vehicle while reducing interplanetary cruise times by more than 3 years [2] Additionally with modern guidance and control, Neptune aerocapture with blunt-body aeroshells is realizable [3][4]. There are limited papers in the literature investigating Uranus aerocapture with those available providing a preliminary feasibility assessment [5]. Consequently, the two-year funded NASA Space Technology Mission Directorate (STMD)-funded project, titled Aerocapture System as an Enabling Technology for Ice Giants Missions, aims to mature the analysis and technology state of Uranus aerocapture. Trajectory Design: This paper presents the current state of the trajectory design in support of the new aerocapture project. The project design philosophy is inspired from recent Neptune aerocapture studies, which employed modern guidance and control, in the sense that blunt-body aeroshells are analyzed. An assessment of the theoretical flight path angle corridor width is conducted for a range of ballistic coefficients and lift-to-drag ratios for both Space Launch System and Falcon Heavy Launch Vehicle interplanetary trajectory solutions. The results from the corridor width assessment provide an assessment of the aerocapture design-space and qualitative metrics on trajectory design considerations. The Program to Optimize Simulated Trajectories II (POST2) is utilized to run Monte Carlo simulations of Uranus aerocapture three-degree-of-freedom bank angle modulated trajectories using a closed-loop numerical-predictor corrector guidance algorithm. UranusGRAM 2021 is utilized as the atmospheric model [6]. A Uranus-developed aerodatabase, originally derived from Mars Science Laboratory (MSL), is utilized to provide vehicle aerodynamics over a wide range of hypersonic flow regimes. A MSL-derived 70 deg 4.5m diameter sphere-cone aeroshell that houses the UOP payload mass is assumed. Robustness testing and performance analysis is conducted, including the assessment of entry state errors, atmosphere density variations, and aerodynamic dispersions. Post-aerocapture Delta-V and aerothermal statistics are formulated into propellant mass and TPS requirements. The results presented in the paper will demonstrate the trajectory viability of Uranus aerocapture. Preliminary Results: Preliminary trajectory design results indicates successful Uranus aerocapture with a blunt-body aeroshell housing the same payload mass as the UOP mission from an Earth-to-Uranus interplanetary trajectory arriving in less than 7 years. From this interplanetary trajectory, aerocapture provides an orbit insertion Delta-V capability of 6.9 km/s requiring less than 300 m/s for post-aerocapture correction burns (15% of wet mass allocated to propellant). To put this into perspective, the UOP study utilized an Earth-to-Uranus trajectory that arrives in 13 years and requires more than 1000 m/s for fully-propulsive orbit insertion (40% of wet mass allocated to propellant). Achieving the same 6.9 km/s Delta-V capability fully-propulsively is mass prohibitive (97% of wet mass allocated to propellant). Nevertheless, aerocapture has the potential to reduce interplanetary transit times to Uranus by half while delivering the same payload mass to orbit in a reduced propellant mass footprint.

Rohan Deshmukh↗

NASA NTRS1976

Design, development, and test of a laser velocimeter for high speed turbomachinery

The paper describes the design, development, and test of a laser Doppler velocimeter (LDV) to map the velocity fields in a small, high speed, 8:1 pressure ratio, 0.91 kg/s (2 lbm/s), centrifugal compressor. This instrument is being used as a diagnostic tool for the study of the basic fluid dynamics of the inducer, impeller and the diffuser regions of high pressure ratio, small, centrifugal compressors. The LDV instrumentation has been optimized to permit the measurement of instantaneous velocities up to approximately 500 m/s, measured in absolute coordinates, within a rotating compressor impeller. Velocities of the same magnitude can also be measured in the two-dimensional radial plane of the diffuser. It is shown that LDV can make nondisturbing gas velocity measurements and surmount the difficulties found in applying other types of aerodynamic instrumentation to this measurement problem.

Runstadler, P. W., Jr.↗

NASA NTRS1976

Some recent progress in transonic flow computation

Although the development of a finite difference relaxation procedure to solve the steady form of equations of motion gave birth to the study of computational transonic aerodynamics and considerable progress has been made using the small disturbance theory, no general analytical solution method yet exists for transonic flows that include three dimensional unsteady, and viscous effects. Two techniques are described which are useful in computational transonic aerodynamics applications. The finite volume method simplifies the application of boundary conditions without introducing the constriction associated with small disturbance theory. Governing equations are solved in a Cartesian coordinate system using a body-oriented and shock-oriented mesh network. Only the volume and surface normal directions of the volume elements must be known. The other method, configuration design by numerical optimization, can be used by aircraft designers to develop configurations that satisfy specific geometric performance constraints. Two examples of airfoil design by numerical optimization are presented.

Ballhaus, W. F.↗

NASA NTRS1991

Applications of low lift to drag ratio aerobrakes using angle of attack variation for control

Several applications of low lift to drag ratio aerobrakes are investigated which use angle of attack variation for control. The applications are: return from geosynchronous or lunar orbit to low Earth orbit; and planetary aerocapture at Earth and Mars. A number of aerobrake design considerations are reviewed. It was found that the flow impingement behind the aerobrake and the aerodynamic heating loads are the primary factors that control the sizing of an aerobrake. The heating loads and other loads, such as maximum acceleration, are determined by the vehicle ballistic coefficient, the atmosphere entry conditions, and the trajectory design. Several formulations for defining an optimum trajectory are reviewed, and the various performance indices that can be used are evaluated. The 'nearly grazing' optimal trajectory was found to provide the best compromise between the often conflicting goals of minimizing the vehicle propulsive requirements and minimizing vehicle loads. The relationship between vehicle and trajectory design is investigated further using the results of numerical simulations of trajectories for each aerobrake application. The data show the sensitivity of the trajectories to several vehicle parameters and atmospheric density variations. The results of the trajectory analysis show that low lift to drag ratio aerobrakes, which use angle of attack variation for control, can potentially be used for a wide range of aerobrake applications.

Mulqueen, J. A.↗