Engineering Papers⌕ Search

SEARCH · Engineering Papers

Results for “EFFECTIVE MASS”

Search indexed NASA NTRS and DOE OSTI research on propulsion, heat transfer, battery materials and energy systems. Follow report and document links to the original sources.

Quote a phrase for an exact phrase match. Source license links do not imply unrestricted reuse.

At least 505 records · Page 28
All sourcesNASA NTRSDOE OSTI
NASA NTRS1999

A Collaborative Analysis Tool for Thermal Protection Systems for Single Stage to Orbit Launch Vehicles

Presented is a design tool and process that connects several disciplines which are needed in the complex and integrated design of high performance reusable single stage to orbit (SSTO) vehicles. Every system is linked to every other system and in the case of SSTO vehicles with air breathing propulsion, which is currently being studied by the National Aeronautics and Space Administration (NASA); the thermal protection system (TPS) is linked directly to almost every major system. The propulsion system pushes the vehicle to velocities on the order of 15 times the speed of sound in the atmosphere before pulling up to go to orbit which results high temperatures on the external surfaces of the vehicle. Thermal protection systems to maintain the structural integrity of the vehicle must be able to mitigate the heat transfer to the structure and be lightweight. Herein lies the interdependency, in that as the vehicle's speed increases, the TPS requirements are increased. And as TPS masses increase the effect on the propulsion system and all other systems is compounded. To adequately determine insulation masses for a vehicle such as the one described above, the aeroheating loads must be calculated and the TPS thicknesses must be calculated for the entire vehicle. To accomplish this an ascent or reentry trajectory is obtained using the computer code Program to Optimize Simulated Trajectories (POST). The trajectory is then used to calculate the convective heat rates on several locations on the vehicles using the Miniature Version of the JA70 Aerodynamic Heating Computer Program (MINIVER). Once the heat rates are defined for each body point on the vehicle, then insulation thicknesses that are required to maintain the vehicle within structural limits are calculated using Systems Improved Numerical Differencing Analyzer (SINDA) models. If the TPS masses are too heavy for the performance of the vehicle the process may be repeated altering the trajectory or some other input to reduce the TPS mass. The problem described is an example of the need for collaborative design and analysis. Analysis tools are being developed to facilitate these collaborative efforts. RECIPE is a cross-platform application capable of hosting a number of engineers and designers across the Internet for distributed and collaborative engineering environments. Such integrated system design environments allow for collaborative team design analysis for performing individual or reduced team studies. The analysis tools mentioned earlier are commonly run on different platforms and are usually run by different people. To facilitate the larger number of potential runs that may need to be made, RECIPE connects the computer codes that calculate the trajectory data, heat rate data, and TPS masses so that the output from each tool is easily transferred to the model input files that need it. This methodology is being applied to solve launch vehicle thermal design problems to shorten the design cycle, and enable the project team to evaluate design options. Results will be presented indicating the effectiveness of this as a collaborative design tool.

Alexander, Reginald A.↗
NASA NTRS1987

Bimodal star formation - Constraints from galaxy colors at high redshift

The possibility that at early epochs the light from elliptical galaxies is dominated by stars with an initial mass function (IMF) which is deficient in low-mass stars, relative to the solar neighborhood is investigated. V-R colors for the optical counterparts of 3CR radio sources offer the most severe constraints on the models. Reasonable fits are obtained to both the blue, high-redshift colors and the redder, low-redshift colors with a model galaxy which forms with initially equal star formation rates in each of two IMF modes: one lacking low-mass stars, and one with stars of all masses. The net effect is that the time-integrated IMF has twice as many high-mass stars as the solar neighborhood IMF, relative to low mass stars. A conventional solar neighborhood IMF does not simultaneously account for both the range in colors at high redshift and the redness of nearby ellipticals, with any single star formation epoch. Models with a standard IMF require half the stellar population to be formed in a burst at low redshift z of about 1.

Wyse, Rosemary F. G.↗
NASA NTRS1998

Influence of Impactor Mass on the Damage Characteristics and Failure Strength of Laminated Composite Plates

The results of an experimental study of the effect of impactor mass on the low-speed impact response of laminated flat composite plates is presented. Dropped weight impact response, damage characteristics, and residual strengths of quasi-isotropic flat plates are presented for a range of energy levels by systematically varying the mass of the impactor. Measured contact forces and damage areas are also presented. The results indicate that the contact force and damage area are nonlinear functions of the impactor mass and vary considerably over the entire range of energy levels considered. The different damage levels induced in a plate specimen when impacted at a given energy level with impactors of different masses significantly influence its compressive residual strength. The results provide clear and consistent trends in contact force, damage area, and compression-after-impact strength when the data are expressed as a function of the impactor momentum.

Ambur, Damodar R.↗
NASA NTRS2000

Non-Tidal Non-Seasonal Oceanic Mass Redistribution Estimated from the TOPEX/Poseidon Observation

Topex/Poseidon altimetry data are used to estimate the non-tidal mass redistribution as a function of space-time. The goal is to study the contribution of ocean circulations in the geodynamic effects including Earth's rotational and gravitational variations. We examine the non-seasonal anomalies at monthly sampling rate over the T/P span of eight years, concentrating especially on interannual variabilities. Since the sea-surface height data obtained from altimetry is the combined effect of steric change (primarily thermal effect) and the mass flux, and because the former has zero contribution to the geodynamic effects, one needs to do a so-called steric correction by removing from the altimetry data the steric contributions. We achieve it using multiyear monthly sea-surface temperature maps, together with monthly "climatology" for mixed-layer depth maps for mean-months of the year. We analyze both sets of the altimetry data and the steric correction data using the empirical orthogonal function/principal component analysis (in which we take care of issues associated with the area-weighting and non-zero mean), and examine the most important modes, either globally or regionally. In particular, the ENSO in the tropical Pacific/Indian Oceans exhibits the most prominent pattern. The net mass transport after the steric correction can then be compared with: (1) ocean general circulation model outputs for the same period of time (such as from POCM-4B); (2) non-atmospheric Earth rotation variations obtained from space geodesy data and atmospheric angular momentum data; (3) non-atmospheric low-degree gravitational variations from satellite-laser-ranging observations and global atmospheric data.

Chao, Benjamin F.↗
NASA NTRS1999

Collaborative Analysis Tool for Thermal Protection Systems for Single Stage to Orbit Launch Vehicles

Presented is a design tool and process that connects several disciplines which are needed in the complex and integrated design of high performance reusable single stage to orbit (SSTO) vehicles. Every system is linked to every other system and in the case of SSTO vehicles with air breathing propulsion, which is currently being studied by the National Aeronautics and Space Administration (NASA); the thermal protection system (TPS) is linked directly to almost every major system. The propulsion system pushes the vehicle to velocities on the order of 15 times the speed of sound in the atmosphere before pulling up to go to orbit which results high temperatures on the external surfaces of the vehicle. Thermal protection systems to maintain the structural integrity of the vehicle must be able to mitigate the heat transfer to the structure and be lightweight. Herein lies the interdependency, in that as the vehicle's speed increases, the TPS requirements are increased. And as TPS masses increase the effect on the propulsion system and all other systems is compounded. To adequately determine insulation masses for a vehicle such as the one described above, the aeroheating loads must be calculated and the TPS thicknesses must be calculated for the entire vehicle. To accomplish this an ascent or reentry trajectory is obtained using the computer code Program to Optimize Simulated Trajectories (POST). The trajectory is then used to calculate the convective heat rates on several locations on the vehicles using the Miniature Version of the JA70 Aerodynamic Heating Computer Program (MINIVER). Once the heat rates are defined for each body point on the vehicle, then insulation thickness that are required to maintain the vehicle within structural limits are calculated using Systems Improved Numerical Differencing Analyzer (SINDA) models. If the TPS masses are too heavy for the performance of the vehicle the process may be repeated altering the trajectory or some other input to reduce the TPS mass.

Alexander, Reginald Andrew↗
NASA NTRS1971

Trapping - A control phenomenon of spinning drag-free satellites.

A drag-free satellite contains a proof mass in an internal cavity and is controlled in translation so that it never touches the proof mass. The satellite shields the proof mass from external forces thus allowing the proof mass to follow a drag-free orbit. Spinning the satellite is desirable because it attenuates the effect of proof mass disturbing forces and simplifies the attitude control. The design of a translation controller for a spinning drag-free satellite typically includes a deadspace to eliminate chatter. This design feature and the inability to locate precisely the mass center give rise to a phenomenon called trapping that potentially could waste significant amounts of propellant. A theory is developed and experimentally verified that explains the role of these factors and provides insight into the effect of other control parameters.

Powell, J. D.↗
NASA NTRS2001

A Collaborative Analysis Tool for Thermal Protection Systems for Single Stage to Orbit Launch Vehicles

Presented is a design tool and process that connects several disciplines which are needed in the complex and integrated design of high performance reusable single stage to orbit (SSTO) vehicles. Every system is linked to all other systems, as is the case with SSTO vehicles with air breathing propulsion, which is currently being studied by the National Aeronautics and Space Administration (NASA). In particular, the thermal protection system (TPS) is linked directly to almost every major system. The propulsion system pushes the vehicle to velocities on the order of 15 times the speed of sound in the atmosphere before pulling up to go to orbit which results in high temperatures on the external surfaces of the vehicle. Thermal protection systems to maintain the structural integrity of the vehicle must be able to mitigate the heat transfer to the structure and be lightweight. Herein lies the interdependency, in that as the vehicle's speed increases, the TPS requirements are increased. And as TPS masses increase the effect on the propulsion system and all other systems is compounded. To adequately calculate the TPS mass of this type of vehicle several engineering disciplines and analytical tools must be used preferably in an environment that data is easily transferred and multiple iterations are easily facilitated.

Alexander, Reginald↗
NASA NTRS1998

Localized infusion of IGF-I results in skeletal muscle hypertrophy in rats

Insulin-like growth factor I (IGF-I) peptide levels have been shown to increase in overloaded skeletal muscles (G. R. Adams and F. Haddad. J. Appl. Physiol. 81: 2509-2516, 1996). In that study, the increase in IGF-I was found to precede measurable increases in muscle protein and was correlated with an increase in muscle DNA content. The present study was undertaken to test the hypothesis that direct IGF-I infusion would result in an increase in muscle DNA as well as in various measurements of muscle size. Either 0.9% saline or nonsystemic doses of IGF-I were infused directly into a non-weight-bearing muscle of rats, the tibialis anterior (TA), via a fenestrated catheter attached to a subcutaneous miniosmotic pump. Saline infusion had no effect on the mass, protein content, or DNA content of TA muscles. Local IGF-I infusion had no effect on body or heart weight. The absolute weight of the infused TA muscles was approximately 9% greater (P < 0.05) than that of the contralateral TA muscles. IGF-I infusion resulted in significant increases in the total protein and DNA content of TA muscles (P < 0.05). As a result of these coordinated changes, the DNA-to-protein ratio of the hypertrophied TA was similar to that of the contralateral muscles. These results suggest that IGF-I may be acting to directly stimulate processes such as protein synthesis and satellite cell proliferation, which result in skeletal muscle hypertrophy.

Non-NASA Center↗
NASA NTRS2010

Expendable Launch Vehicles Briefing and Basic Rocketry Physics

This slide presentation is composed of two parts. The first part shows pictures of launch vehicles and lift offs or in the case of the Pegasus launch vehicle separations. The second part discusses the basic physics of rocketry, starting with Newton's three physical laws that form the basis for classical mechanics. It includes a review of the basic equations that define the physics of rocket science, such as total impulse, specific impulse, effective exhaust velocity, mass ratio, propellant mass fraction, and the equations that combine to arrive at the thrust of the rocket. The effect of atmospheric pressure is reviewed, as is the effect of propellant mix on specific impulse.

Delgado, Luis G.↗
NASA NTRSDate not supplied

A Multi-Disciplinary Analysis Framework for the Design of Small Launch Vehicles

Between the years of 1995 and 2014 the number of small satellites (1-500 kg) went up from 20 to 180. [1] Out of the 180 launched in 2014 66% were Nano satellites (1-10 kg). [1] With this trend of smaller satellites, one would expect a rise in number of small launch vehicles (SLVs are defined by capability to carry 1-100 kg to orbit), but this has not happened: [3] only 8% of all small satellites are launched on SLVs and the others become secondary payloads on regular launch vehicles. [1] This results in small satellites being placed into either suboptimal orbits or waiting for launch schedules to align with bigger launches, resulting in long waiting times. Dedicated SLVs could improve the responsiveness of small satellite launches, but SLV design is complicated by the large architecture space needed to be explored in order to find efficient and affordable designs. The SLV architecture trade space contains many discrete options, e.g. solid vs. liquid fuels, air launch vs. ground launch, number of stages, etc. [2][3] Performing detailed analysis for all the options at once would be prohibitive. Thus, a sizing environment/framework that is capable of providing necessary information for conceptual-level trade studies and can rapidly explore the vast SLV architecture space is necessary. The framework, illustrated in Figure 1, consists of four disciplines central to the sizing and synthesis of launch vehicles: propulsion, aerodynamics, structures, and trajectory. For the aerodynamics and trajectory disciplines, Missile DATCOM and POST2 are used, respectively. The propulsion and structures disciplines are represented in the framework with tools developed at ASDL Georgia Tech. For propulsion, the Solid Motor Analysis Code (SMAC) is a physics-based conceptual design tool for solid rocket motors. SMAC is capable of geometric burn simulation, ballistic analysis, and prediction of thrust performance. [4] For structures, Launch Vehicle Structural Analysis (LVSA) tool is a physics-based tool that focuses on structural dynamic analysis with sizing capability. These tools are integrated into the framework illustrated in Figure 1 with the corresponding connections described in Table 1. The process flow is as follows: first, SMAC sizes insulation and calculates maximum operating pressures for each vehicle stage. The MEOP and insulation thickness values from SMAC are fed to LVSA which then utilizes this information to size the motor casing. This creates a feedback loop between LVSA and SMAC that converges on the radius available for fuel, casing, and insulation thickness. Once the stage sizing is converged on SMAC creates an engine deck that is passed to POST2. Next, the data from SMAC and LVSA goes into Missile DATCOM which generates an aerodynamics database for POST2. Finally, POST2 performs a targeting optimization while maximizing the payload mass to orbit. Within the framework, POST2 is automated in order to be robust to a wide variety of possible designs by performing a Monte Carlo simulation over the initialization vector of the POST2 optimization variables. To demonstrate the capability of this framework, a sample problem of exploring the design space of an SLV capable of placing satellites into a low Earth orbit (inclination=47 deg, 196.5 by 211.3 nm) is used. This sample problem is a ground-launched SLV consisting of four in-line SRM stages. The multidisciplinary design analysis (MDA) environment is utilized to explore a design space consisting of 22 continuous and 12 discrete variables, shown by Table 1 by blocks 1,2,3. Running a full factorial design of experiments (DOE) would have been prohibitively expensive even with this reduced design space, thus a space filling design with 3,502 and then additional expansion of 2,602 cases was used. The first DOE consisted of 3,502 cases, and all of the variables were varied. These input variables are listed in blocks 1, 2, and 3 in Table 1. Most of the variables are propulsion related with stage lengths determining the delta-V split of the SLV. The expansion consisted of 2,602 cases, and the continuous variables were set to be equal to the most promising designs from the sizing of first DOE. For each set of continuous variables, the discrete variables (grain type, star points, and propellant) were to varied. The results of the DOE can be seen in Figure 2; each of the points in this plot represents a closed launch vehicle that reaches the targeted orbit. For each of the cases, there is information on flown trajectory, structural, and propulsion properties of the SLV. For example, Figure 3 shows changes in altitude and velocity with time for a particular case. The right side of Figure 3 clearly shows the coasting (slow decrease) and burning phases (sharp increase) of the SLV mission. LV mass is positively correlated with the optimized payload mass to orbit because heavier LVs carry more fuel and thus have more stored chemical energy. Furthermore, for any given payload mass to orbit, the most efficient design would result in the smallest LV. The results as visualized in Figure 2 shows this tradeoff, and the Pareto frontier of the efficient designs can be seen along the dotted line. Figure 2 can be divided into regions with the lowest mass vehicles corresponding to the minimum bound on radius, and the highest mass vehicles corresponding to the maximum bound on radius. Within a mass region, the discrete variables, such as propellant type and propellant grain arrangement, have the most effect on payload mass. This paper presents an MDA framework that can perform an automated physics-based sizing of SLV designs and a corresponding methodology to utilize the MDA to explore the design space of SLVs. The proposed methodology was applied to a perform a design space exploration for a four stage SLV. The outputs show the expected pareto frontier forming and provide detailed information about the SLV performance and staging. Using the produced data, it will be possible to select a set of pareto optimal designs that can then be further explored in subsequent design cycles. This demonstration shows that automated design space exploration should be used in the early phase design of future SLV concepts.

Nikita S Birbasov↗
NASA NTRS1987

Nonlinear simulation of a flexible aircraft in maneuvering flight

A mathematical model integrating nonlinear rigid-body flight mechanics and linear aeroelastic dynamics is examined; the equations of motion for an elastic aircraft in accelerated flight are developed using Lagrangian mechanics. This approach is used to construct a simulation model of an F/A-18 (configured with tip missiles) which includes angular/elastic inertial coupling (IC). In general, the elastic modes significantly affected by IC were aerodynamically decoupled from the rest of the model. The affected modes were those which induced changes in total aircraft mass distribution. The elastic effect is noticeable if deformation-induced mass distribution changes are significant with respect to modal mass and modal frequencies. A modal parameter is presented which characterizes the level of IC between elastic momentum and rigid-body angular momentum.

Buttrill, Carey S.↗
NASA NTRS1997

Application of Shear Plate Interferometry to Jet Diffusion Flame Temperature Measurements

The recent ban on the production of bromotrifluoromethane (CF3Br) because of its high stratospheric ozone depletion potential has led to interest in finding alternative agents for fire extinguishing applications. Some of the promising alternatives are fluorinated hydrocarbons. A clear understanding of the effects of CF3Br and alternative chemical suppressants on diffusion flames is therefore necessary in the selection of alternative suppressants for use in normal and microgravity. The flame inhibition effects of halogen compounds have been studied extensively in premixed systems. The effect of addition of halocarbons (carbon-halogen compounds) to diffusion flames has been studied experimentally in coflow configurations and in counterflow gaseous and liquid-pool flames. Halogenated compounds are believed to inhibit combustion by scavenging hydrogen radicals to form the relatively unreactive compound HF, or through a catalytic recombination cycle involving HBr to form H2. Comparisons between halogens show that bromine inhibition is significantly more effective than chlorine or fluorine. Although fluorinated compounds are only slightly more effective inhibitors on a mass basis than nitrogen, they are more effective on a volume basis and are easily stored in liquid form. The objectives of this study are (a) to determine the stability limits of laminar jet diffusion flames with respect to inhibitor concentration in both normal and microgravity, and (b) to investigate the structure of halocarbon-inhibited flames. In the initial phase of this project, visual diagnostics were used to observe the structure and behavior of normal and microgravity flames. The initial observations showed significant changes in the structure of the flames with the addition of halocarbons to the surrounding environment, as discussed below. Furthermore, the study established that the flames are more stable relative to the addition of halocarbons in microgravity than in normal gravity. Visual diagnostics of flames are, however, necessarily limited to detection of radiative emission in the visible range, and offer only qualitative information about the nature of the processes in the flame. In particular, the study sought to understand the structure of the inhibitor-perturbed flames with regard to temperature and species concentration in the outer region of the flame. Whereas thermocouple measurements can be used in ground based studies, their implementation in drop-tower rigs is limited. A possible approach to determine the temperature field around the flame is to use interferometric techniques. The implementation and testing of a shear-plate interferometry technique is described below.

VanDerWege, Brad A.↗
NASA NTRS2019

Observations of Surface Mass Balance on Pine Island Glacier, West Antarctica, and the Effect of Strain History in Fast-Flowing Sections

Surface mass balance (SMB) is the net input of mass on a glacier's upper surface, composed of snow deposition, melt and erosion processes, and is a major contributor to the overall mass balance. Pine Island Glacier (PIG) in West Antarctica has been dynamically imbalanced since the early 1990s, indicating that discharge of solid ice into the oceans exceeds snow deposition. However, observations of the SMB pattern on the fast flowing regions are scarce, and are potentially affected by the firn's strain history. Here, we present new observations from radar-derived stratigraphy and a relatively dense network of firn cores, collected along a ~900 km traverse of PIG. Between 1986 and 2014, the SMB along the traverse was 0.505 m w.e. a−1 on average with a gradient of higher snow deposition in the South-West compared with the North-East of the catchment. We show that along ~80% of the traverse the strain history amounts to a misestimation of SMB below the nominal uncertainty, but can exceed it by a factor 5 in places, making it a significant correction to the SMB estimate locally. We find that the strain correction changes the basin-wide SMB by ~0.7 Gt a−1 and thus forms a negligible (1%) correction to the glacier's total SMB.

ice core↗
NASA NTRS2004

Effects of Heat Flux, Oxygen Concentration and Glass Fiber Volume Fraction on Pyrolysate Mass Flux from Composite Solids

Experimental work on the effects of heat flux, oxygen concentration and glass fiber volume fraction on pyrolysate mass flux from samples of polypropylene/glass fiber composite (PP/G) is underway. The research is conducted as part of a larger project to develop a test methodology for flammability of materials, particularly composites, in the microgravity and variable oxygen concentration environment of spacecraft and space structures. Samples of PP/G sized at 30x30x10 mm are flush mounted in a flow tunnel, which provides a flow of oxidizer over the surface of the samples at a fixed value of 1 m/s and oxygen concentrations varying between 18 and 30%. Each sample is exposed to a constant external radiant heat flux at a given value, which varies between tests from 10 to 24 kW/m2. Continuous sample mass loss and surface temperature measurements are recorded for each test. Some tests are conducted with an igniter and some are not. In the former case, the research goal is to quantify the critical mass flux at ignition for the various environmental and material conditions described above. The later case generates a wider range of mass flux rates than those seen prior to ignition, providing an opportunity to examine the protective effects of blowing on oxidative pyrolysis and heating of the surface. Graphs of surface temperature and sample mass loss vs. time for samples of 30% PPG at oxygen concentrations of 18 and 21% are presented in the figures below. These figures give a clear indication of the lower pyrolysis rate and extended time to ignition that accompany a lower oxygen concentration. Analysis of the mass flux rate at the time of ignition gives good repeatability but requires further work to provide a clear indication of mass flux trends accompanying changes in environmental and material properties.

Rich, D. B.↗
NASA NTRS2004

Effects of Heat Flux, Oxygen Concentration and Glass Fiber Volume Fraction on Pyrolysate Mass Flux from Composite Solids

Experimental work on the effects of heat flux, oxygen concentration and glass fiber volume fraction on pyrolysate mass flux from samples of polypropylene/glass fiber composite (PP/G) is underway. The research is conducted as part of a larger project to develop a test methodology for flammability of materials, particularly composites, in the microgravity and variable oxygen concentration environment of spacecraft and space structures. Samples of PP/G sized at 30 x 30 x 10 mm are flush mounted in a flow tunnel, which provides a flow of oxidizer over the surface of the samples at a fixed value of 1 m/s and oxygen concentrations varying between 18 and 30%. Each sample is exposed to a constant external radiant heat flux at a given value, which varies between tests from 10 to 24 kW/sq m. Continuous sample mass loss and surface temperature measurements are recorded for each test. Some tests are conducted with an igniter and some are not. In the former case, the research goal is to quantify the critical mass flux at ignition for the various environmental and material conditions described above. The later case generates a wider range of mass flux rates than those seen prior to ignition, providing an opportunity to examine the protective effects of blowing on oxidative pyrolysis and heating of the surface. Graphs of surface temperature and sample mass loss vs. time for samples of 30% PPG at oxygen concentrations of 18 and 21% are presented in the figures below. These figures give a clear indication of the lower pyrolysis rate and extended time to ignition that accompany a lower oxygen concentration. Analysis of the mass flux rate at the time of ignition gives good repeatability but requires further work to provide a clear indication of mass flux trends accompanying changes in environmental and material properties.

Rich, D. B.↗
NASA NTRS1982

Optimal preparation of the ECC ozonesonde

The ECC background current was identified as the removal of residual tri-iodide (iodine) as the cell approaches equilibrium. The altitude dependence of this source of the background current is expected to be only slowly changed in the troposphere with a more rapid decrease in the stratosphere. Oxygen does not play a role in the background current except in the unlikely situation where the electrodes have had all forms of iodine removed from them and the electrodes have not re-equilibrated with the sonde solutions before use. A solution mass transport parameter in the ECC was identified and its altitude dependence determined. The mass transport of tri-iodide dominates in the chemical transduction of ozone to electrical signal. The effect of the mass transport on the ECC background current is predicted. An electrochemical model of the ECC has been developed to predict the response of the ECC to various ozone vertical profiles. The model corresponds very closely to the performance of the ECC in the laboratory. Based on this model, an ECC with no background current is predicted to give total ozone values within 1% of the correct value, although the vertical profile may be in error by as much as + or - 15%.

Thornton, D. C.↗