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A discussion is given of the ongoing research related to laminar flow airfoils, nacelles, and wings where the laminar flow is maintained by a favorable pressure gradient, surface suction or a combination of the two. Design methologies for natural laminar flow airfoil sections and wings for both low and high speed applications are outlined. Tests of a 7-foot chord, 23° sweep laminar-flow-control-airfoil at high subsonic Mach numbers are described along with the associated stability theory used to design the suction system. The state-of-the-art of stability theory is simply stated and a typical calculation illustrated. In addition recent computer simulations of transition using the time dependent Navier-Stokes (N-S) equations are briefly described. Advances in wind tunnel capabilities and instrumentation will be reviewed followed by the presentation of a few results from both wind tunnels and flight. Finally, some suggestions for future work will complete the paper.

Solar proton events (SPEs) represent the single-most significant source of acute radiation exposure during space missions. Historically, an exponential in rigidity (particle momentum) fit has been used to express the SPE energy spectrum using GOES data up to 100 MeV. More recently, researchers have found that a Weibull fit better represents the energy spectrum up to 1000 MeV (1 GeV). In addition, the availability of SPE data extending up to several GeV has been incorporated in analyses to obtain a more complete and accurate energy spectrum representation. In this paper we discuss the major SPEs that have occurred over the past five solar cycles (~50+ years) in detail - in particular, Aug 1972 and Sept & Oct 1989 SPEs. Using a high-energy particle transport/dose code, radiation exposure estimates are presented for various thicknesses of aluminum. The effects on humans and spacecraft systems are also discussed in detail.

A new method is developed for calculating the low energy neutron flux in a space environment which is protected from galactic cosmic rays (GCR) and solar particle events (SPE) by shielding materials. Our calculations are compared with low energy neutron flux flight data recorded on four different STS low earth orbit missions. We also compare our neutron flux calculations with the low energy neutron flux data recorded by MIR. The low energy neutron flux calculations can be described as a deterministic method for solving the Boltzmann equation for the light ion flux associated with a given environment. Existing Monte Carlo neutron flux simulations associated with the MIR and ISS space stations are also compared with our deterministic method for calculating neutron flux.

Many studies have been performed to quantify the formation and evolution of roughness on ice shapes created in Appendix C icing conditions, which exhibits supercooled liquid droplets ranging from 1-50 µm. For example Anderson and Shin (1997), Anderson et al. (1998), and Shin (1994) represent early studies of ice roughness during short-duration icing events measured in the Icing Research Tunnel at the NASA Glenn Research Center. In the historical literature, image analysis techniques were employed to characterize the roughness. Using multiple images of the roughness elements, these studies of roughness focused on extracting parametric representations of ice roughness elements. While the image analysis approach enabled many insights into icing physics, recent improvements in laser scanning approaches have revolutionized the process of ice accretion shape characterization.

Reynolds number effects noted from selected test programs conducted in the Langiey 0.3-Meter Transonic Cryogenic Tunnel (0.3-m TCT) are discussed. The tests, which cover a unit Reynolds number range from about 2.0 to 80.0 million per foot, summarize effects of Reynolds number on: 1) aerodynamic data from a supercritical airfoil, 2) results from several wall interference correction techniques, and 3) results obtained from advanced, cryogenic test techniques. The test techniques include 1) use of a cryogenic sidewall boundary layer removal system, 2) detailed pressure and hot wire measurements to determine test section flow quality, and 3) use of a new hot film system suitable for transition detection in a cryogenic wind tunnel. The results indicate that Reynolds number effects appear most significant when boundary layer transition effects are present and at high lift conditions when boundary layer separation exists on both the model and the tunnel sidewall.

The direct first order effect of Reynolds number on the determination of drag-divergence conditions is summarized for six airfoils which were tested in the Langley 0.3-Meter Transonic Cryogenic Tunnel. A second order effect, derived through the effect of Reynolds number on the sidewall boundary layer, is included. In addition, a comparison of how the drag-divergence condition is affected on going from one class of airfoil to another is presented. The drag-divergence condition is affected first order by Reynolds number for each of the six airfoils and of course all data are affected second order, since the presence of the boundary layer necessitates a sidewall correction.

In view of manned missions targeted to the Moon, for which radiation exposure is one of the greatest challenges to be tackled, it is of fundamental importance to have available a tool, which allows determination of the particle flux and spectra at any time and at any point of the lunar surface. With this goal in mind, a new model of the Moon’s radiation environment due to Galactic Cosmic Rays (GCR) and Solar Particle Events (SPE) has been developed. Primary particles reach the lunar surface, and are transported all throughout the subsurface layers, with backscattering patterns taken into account. The surface itself has been modeled as regolith and bedrock, with composition taken from the results of the instruments flown on the Apollo missions, namely on the Apollo 12 from the Oceanus Procellarum landing site. Subsurface environments like lava tubes have been considered in the analysis.

This paper discusses a project for adapting advanced technology, much of it borrowed from the jet transport, to general aviation design practice. The NASA funded portion of the work began in 1969 at the University of Kansas and resulted in a smaller, experimental wing with spoilers and powerful flap systems for a Cessna Cardinal airplane. The objective was to obtain increased cruise performance and improved ride quality while maintaining the take-off and landing speeds of the unmodified airplane. Some flight data and research pilot comments are presented. The project was expanded in 1972 to include a light twin-engine airplane. For the twin there was the added incentive of a potential increase in single-engine climb performance. The expanded project is a joint effort involving the University of Kansas, Piper Aircraft Company, Robertson Aircraft Company, and Wichita State University. The use of a new high-lift Whitcomb airfoil is planned for both the wing and the propellers.

A research program was conducted to evaluate the effectiveness of icing tunnel hybrid model design. A hybrid design is where the full-scale leading edge of a wing section is maintained only to a certain percentage of the local chord, while the aft section of the model is redesigned into a shortened or truncated planform. An initial study was conducted in 2020 where the ice shape geometries on a full-chord length version of the swept CRM65 wing model were compared to those from the hybrid version of CRM65 that were obtained in the NASA Icing Research Tunnel in 2015. The results were reported in a 2021 paper. For most test conditions, the overall size and shape of the ice shapes compared well. However, the ice shapes from the full-chord model were generally slightly smaller than those from the hybrid model.

Artificial ice shapes of various geometric fidelity were tested on a wing model based on the Common Research Model. Low Reynolds number tests were conducted at Wichita State University’s Walter H. Beech Memorial Wind Tunnel utilizing an 8.9% scale model, and high Reynolds number tests were conducted at ONERA’s F1 wind tunnel utilizing a 13.3% scale model. Several identical geometrically-scaled ice shapes were tested at both facilities, and the results were compared at overlapping Reynolds and Mach numbers. This was to ensure that the results and trends observed at low Reynolds number could be applied and continued to high, near-flight Reynolds number. The data from Wichita State University and ONERA F1 agreed well at matched Reynolds and Mach numbers. The lift and pitching moment curves agreed very well for most configurations.

In-flight icing is an important consideration that affects aircraft design, performance, certification and safety. Newer regulations combined with increasing demand to reduce fuel burn, emissions and noise are driving a need for improvements in icing simulation capability. To that end, this paper presents the results of additional ice accretion testing conducted in the NASA Icing Research Tunnel in January 2022 with a large swept wing section typical of a modern commercial transport. The model was based upon a section of the Common Research Model wing at the 64% semispan station with a streamwise chord length of 136 in. The test conditions were developed with an icing scaling analysis to generate similar conditions for a small median volumetric diameter (MVD) = 25 μm cloud and a large MVD = 110 μm cloud. A series of tests were conducted over a range of total temperature from -23.8 °C to -1.4 °C with all other conditions held constant.

An advanced analytical methodology has been developed for predicting the residual strength of stiffened thin-sheet riveted shell structures such as those used for the fuselage of a commercial transport aircraft. The crack-tip opening angle elastic-plastic fracture criterion has been coupled to a geometric and material nonlinear finite element shell code for analyzing complex structural behavior. An automated adaptive mesh refinement capability together with global-local analysis methods have been developed to predict the behavior of fuselage structure with long cracks. This methodology is currently being experimentally verified. Advanced nondestructive inspection technology has been developed that will provide airline operators with the capability to conduct reliable and economical broad-area inspections of aircraft structures.

Design data are presented for a class of high-performance single-engine business airplanes. The design objectives include a cruise speed of 300 knots, a cruise altitude of 10,700 m (35,000 ft), a cruise payload of six passengers (including crew and baggage), and a no-reserves cruise range of 1300 n.mi. Two unconventional aerodynamic technologies were evaluated: the individual and combined effects of cruise-matched wing loading and of a natural laminar flow airfoil were analyzed. The tradeoff data presented illustrate the ranges of wing geometries, propulsion requirements, airplane weights, and aerodynamic characteristics which are necessary to meet the design objectives. very large design and performance improvements resulted from use of the aerodynamic technologies evaluated. Is is shown that the potential exists for achieving more than 200-percent greater fuel efficiency than is achieved by current airplanes capable of similar cruise speeds, payloads, and ranges.

An experimental research effort was begun to develop a database of airplane aerodynamic characteristics with simulated ice accretion over a large range of incidence and sideslip angles. Wind-tunnel testing was performed at the NASA Langley 12-ft Low-Speed Wind Tunnel using a 3.5% scale model of the NASA Langley Generic Transport Model. Aerodynamic data were acquired from a six-component force and moment balance in static-model sweeps from α = -5 to 85 deg. and β = -45 to 45 deg. at a Reynolds number of 0.24x10⁶ and Mach number of 0.06. The 3.5% scale GTM was tested in both the clean configuration and with full-span artificial ice shapes attached to the leading edges of the wing, horizontal and vertical tail. Aerodynamic results for the clean airplane configuration compared favorably with similar experiments carried out on a 5.5% scale GTM.

This paper describes recent research in integrated aerodynamic-performance design optimization applied to a supersonic transport wing. The subsonic and supersonic aerodynamics are modeled with linear theory and the aircraft performance is evaluated by using a complete mission analysis. The goal of the optimization problem is to either maximize the aircraft range or minimize the take-off gross weight while constraining the total fuel load and approach speed. A major difficulty encountered during this study was the inability to obtain accurate derivatives of the aerodynamic models with respect to the planform shape. This work addresses this problem and provides one solution for the derivative difficulties. Additional optimization studies reveal the impact of camber design on the global optimization problem. In these studies, the plan-form optimization is first conducted on a flat plate wing and camber optimization is performed on the resulting planform.

Fuel costs comprise a major portion of air transport operating costs. Thus, energy efficiency is an essential design goal for future transport aircraft. Advanced composite structures, advanced wing geometries, and active control systems all promise substantial benefits in fuel efficiency and direct operating cost for derivative and new aircraft introduced by 1985. Technology for maintenance of a laminar boundary layer in cruise offers great benefits in fuel efficiency and direct operating cost and may be ready for application to transports introduced in the 1990's. NASA and the air transport industry are cooperating in a comprehensive Aircraft Energy Efficiency Program to expedite the introduction of these advanced technologies into production aircraft.

Alternate Environmental Control and Life Support System (ECLSS) technologies were evaluated to reduce Space Station resources and dependence on expendables resupplied from Earth to sustain a multiperson crew in low-Earth orbit. Options were evaluated to close the oxygen (O2) loop by removing carbon dioxide (CO2) from the cabin air, reducing the CO2 to water, and electrolyzing the water to provide metabolic O2 for crew consumption. Options were also evaluated to close the urine/flush, condensate, and hygiene water loops to provide potable water for crew use. Specific evaluation parameters were derived which included weight, power, volume, maintenance, resupply consumables, and technology readiness.

This paper discusses technologies and software tools that are being implemented in a software toolkit currently under development at NASA Glenn Research Center. Its purpose is to help study the effects of icing on airfoil performance and assist with the aerodynamic simulation process which consists of characterization and modeling of ice geometry, application of block topology and grid generation, and flow simulation. Tools and technologies for each task have been carefully chosen based on their contribution to the overall process. For the geometry characterization and modeling, we have chosen an interactive rather than automatic process in order to handle numerous ice shapes. An Appendix presents features of a software toolkit developed to support the interactive process. Approaches taken for the generation of block topology and grids, and flow simulation, though not yet implemented in the software, are discussed with reasons for why particular methods are chosen.

This paper presents experimental methods for investigating large droplet impingement dynamics and for obtaining small and large water droplet impingement data. Droplet impingement visualization experiments conducted in the Goodrich Icing Wind Tunnel with a 21-in chord NACA 0012 airfoil demonstrated considerable droplet splashing during impingement. The tests were performed for speeds in the range 50 to 175 mph and with cloud median volumetric diameters in the range of 11 to 270 microns. Extensive large droplet impingement tests were conducted at the NASA Glenn Icing Research Tunnel (IRT). Impingement data were obtained for a range of airfoil sections including three 36-inch chord airfoils (MS(1)-0317, GLC-305, and NACA 652-415), a 57-inch chord Twin Otter horizontal tail section and 22.5-minute and 45-minute LEWICE glaze ice shapes for the Twin Otter tail section. Small droplet impingement tests were also conducted for selected test models.

A new Green’s function code (GRNTRN) capable of simulating HZE ions with either laboratory or space boundary conditions is currently under development. The computational model consists of combinations of physical perturbation expansions based on the scales of atomic interaction, multiple scattering, and nuclear reactive processes with use of the Neumann-asymptotic expansions with non-perturbative corrections. The code contains energy loss due to straggling, nuclear attenuation, nuclear fragmentation with energy dispersion and downshifts. Recent publications have focused on code validation in the laboratory environment and have shown that the code predicts energy loss spectra accurately as measured by solid-state detectors in ion beam experiments. In this paper emphasis is placed on code validation with space boundary conditions.