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Gong, L.

Publications and source records attributed to Gong, L..

12 records
All sourcesNASA NTRSDOE OSTI
DOE OSTI2024

Extracting the femtometer structure of strange baryons using the vacuum polarization effect

Abstract One of the fundamental goals of particle physics is to gain a microscopic understanding of the strong interaction. Electromagnetic form factors quantify the structure of hadrons in terms of charge and magnetization distributions. While the nucleon structure has been investigated extensively, data on hyperons are still scarce. It has recently been demonstrated that electron-positron annihilations into hyperon-antihyperon pairs provide a powerful tool to investigate their inner structure. We present a method useful for hyperon-antihyperon pairs of different types which exploits the cross section enhancement due to the effect of vacuum polarization at theJ/ψresonance. Using the 10 billionJ/ψevents collected with the BESIII detector, this allows a precise determination of the hyperon structure function. The result is essentially a precise snapshot of the$$\bar{\Lambda }{\Sigma }^{0}\,(\Lambda {\bar{\Sigma }}^{0})$$ Λ ¯ Σ 0 ( Λ Σ ¯ 0 ) transition process, encoded in the transition form factor ratio and phase. Their values are measured to beR = 0.860 ± 0.029(stat.) ± 0.015(syst.),$$\Delta {\Phi }_{\bar{\Lambda }{\Sigma }^{0}}=(1.011\pm 0.094({{\rm{stat.}}})\pm 0.010({{\rm{syst.}}}))\,{{\rm{r}}}ad$$ Δ Φ Λ ¯ Σ 0 = ( 1.011 ± 0.094 ( stat. ) ± 0.010 ( syst. ) ) r a d and$$\Delta {\Phi }_{\Lambda {\bar{\Sigma }}^{0}}=(2.128\pm 0.094({{\rm{stat.}}})\pm 0.010({{\rm{syst.}}}))\,{{\rm{r}}}ad$$ Δ Φ Λ Σ ¯ 0 = ( 2.128 ± 0.094 ( stat. ) ± 0.010 ( syst. ) ) r a d . Furthermore, charge-parity (CP) breaking is investigated in this reaction and found to be consistent with CP symmetry.

Science & Technology - Other Topics↗
DOE OSTI2024

Search for the leptonic decays D * + → e + ν e and D * + → μ + ν μ

We present the first search for the leptonic decays D * + → e + ν e and D * + → μ + ν μ by analyzing a data sample of electron-positron collisions recorded with the BESIII detector at center-of-mass energies between 4.178 and 4.226 GeV, corresponding to an integrated luminosity of 6.32 fb − 1 . No significant signal is observed. The upper limits on the branching fractions for D * + → e + ν e and D * + → μ + ν μ are set to be 1.1 × 10 − 5 and 4.3 × 10 − 6 at 90% confidence level, respectively. Published by the American Physical Society 2024

Astronomy & Astrophysics↗
DOE OSTI2024

Measurement of Born cross section of $$ {e}^{+}{e}^{-}\to {\Sigma}^{+}{\overline{\Sigma}}^{-} $$ at center-of-mass energies between 3.510 and 4.951 GeV

Abstract Using 24.1 fb −1 ofe + e − collision data collected with the BESIII detector at the BEPCII collider, the Born cross sections and effective form factors of the$$ {e}^{+}{e}^{-}\to {\Sigma}^{+}{\overline{\Sigma}}^{-} $$ e + e − → Σ + Σ ¯ − reaction are measured. The measurements are performed at center-of-mass energies ranging from 3.510 to 4.951 GeV. No significant evidence for the decay of the charmonium(-like) states,ψ(3770),ψ(4040),ψ(4160),Y(4230),Y(4360),ψ(4415), andY(4660), into a$$ {\Sigma}^{+}{\overline{\Sigma}}^{-} $$ Σ + Σ ¯ − final state is observed. Consequently, upper limits for the products of the branching fractions and the electronic partial widths at the 90% confidence level are reported for these decays.

Physics↗
DOE OSTI2024

Measurement of absolute branching fractions of $$ {\textrm{D}}_{\textrm{s}}^{+} $$ hadronic decays

Abstract Usinge + e − collision data collected at the BESIII detector at center-of-mass energies between 4.128 and 4.226 GeV, corresponding to an integrated luminosity of 7.33 fb −1 , we determine the absolute branching fractions of fifteen hadronic$$ {D}_s^{+} $$ D s + decays with a double-tag technique. In particular, we make precise measurements of the branching fractions$$ \mathcal{B}\left({D}_s^{+}\to {K}^{+}{K}^{-}{\pi}^{+}\right)=\left(5.49\pm 0.04\pm 0.07\right)\% $$ B D s + → K + K − π + = 5.49 ± 0.04 ± 0.07 % ,$$ \mathcal{B}\left({D}_s^{+}\to {K}_S^0{K}^{+}\right)=\left(1.50\pm 0.01\pm 0.01\right)\% $$ B D s + → K S 0 K + = 1.50 ± 0.01 ± 0.01 % and$$ \mathcal{B}\left({D}_s^{+}\to {K}^{+}{K}^{-}{\pi}^{+}{\pi}^0\right)=\left(5.50\pm 0.05\pm 0.11\right)\% $$ B D s + → K + K − π + π 0 = 5.50 ± 0.05 ± 0.11 % , where the first uncertainties are statistical and the second ones are systematic. TheCPasymmetries in these decays are also measured and all are found to be compatible with zero.

Physics↗
NASA NTRS1984

Thermal response of Space Shuttle wing during reentry heating

A structural performance and resizing (SPAR) finite element thermal analysis computer program was used in the heat transfer analysis of the Space Shuttle Orbiter that was subjected to reentry aerodynamic heatings. One wing segment of the right wing (WS 240) and the whole left wing were selected for the thermal analysis. Results showed that the predicted thermal protection system (TPS) temperatures were in good agreement with the space transportation system, trajectory 5 (STS-5) flight-measured temperatures. In addition, calculated aluminum structural temperatures were in fairly good agreement with flight data up to the point of touchdown. Results also showed that the internal free convection has a considerable effect on the change of structural temperatures after touchdown.

Gong, L.↗
NASA NTRS1984

Thermal response of Space Shuttle wing during reentry heating

A structural performance and resizing (SPAR) finite element thermal analysis computer program was used in the heat transfer analysis of the space shuttle orbiter that was subjected to reentry aerodynamic heatings. One wing segment of the right wing (WS 240) and the whole left wing were selected for the thermal analysis. Results showed that the predicted thermal protection system (TPS) temperatures were in good agreement with the space transportation system, trajectory 5 (STS-5) flight-measured temperatures. In addition, calculated aluminum structural temperatures were in fairly good agreement with the flight data up to the point of touchdown. Results also showed that the internal free convection had a considerable effect on the change of structural temperatures after touchdown.

Gong, L.↗
NASA NTRS1982

Reentry heating analysis of space shuttle with comparison of flight data

Surface heating rates and surface temperatures for a space shuttle reentry profile were calculated for two wing cross sections and one fuselage cross section. Heating rates and temperatures at 12 locations on the wing and 6 locations on the fuselage are presented. The heating on the lower wing was most severe, with peak temperatures reaching values of 1240 C for turbulent flow and 900 C for laminar flow. For the fuselage, the most severe heating occured on the lower glove surface where peak temperatures of 910 C and 700 C were calculated for turbulent flow and laminar flow, respectively. Aluminum structural temperatures were calculated using a finite difference thermal analyzer computer program, and the predicted temperatures are compared to measured flight data. Skin temperatures measured on the lower surface of the wing and bay 1 of the upper surface of the wing agreed best with temperatures calculated assuming laminar flow. The measured temperatures at bays two and four on the upper surface of the wing were in quite good agreement with the temperatures calculated assuming separated flow. The measured temperatures on the lower forward spar cap of bay four were in good agreement with values predicted assuming laminar flow.

Gong, L.↗
NASA NTRS1982

Reentry heat transfer analysis of the space shuttle orbiter

A structural performance and resizing finite element thermal analysis computer program was used in the reentry heat transfer analysis of the space shuttle. Two typical wing cross sections and a midfuselage cross section were selected for the analysis. The surface heat inputs to the thermal models were obtained from aerodynamic heating analyses, which assumed a purely turbulent boundary layer, a purely laminar boundary layer, separated flow, and transition from laminar to turbulent flow. The effect of internal radiation was found to be quite significant. With the effect of the internal radiation considered, the wing lower skin temperature became about 39 C (70 F) lower. The results were compared with fight data for space transportation system, trajectory 1. The calculated and measured temperatures compared well for the wing if laminar flow was assumed for the lower surface and bay one upper surface and if separated flow was assumed for the upper surfaces of bays other than bay one. For the fuselage, good agreement between the calculated and measured data was obtained if laminar flow was assumed for the bottom surface. The structural temperatures were found to reach their peak values shortly before touchdown. In addition, the finite element solutions were compared with those obtained from the conventional finite difference solutions.

Ko, W. L.↗
NASA NTRS1981

Preflight reentry heat transfer analysis of Space Shuttle

Preflight predictions of the structural temperature distributions during entry are compared with data from the initial Shuttle flight. Finite element thermal analysis programming was used to model the heat flow on Shuttle structures and actual gas properties of air were employed in the analyses of aerodynamic heating. Laminar, separated, and turbulent heat fluxes were calculated for varying locations on the craft using velocity-attitude and angle-of-attack projections taken from the nominal STS-1 trajectory. Temperature time histories of the first flight are compared with laminar and turbulent flow assumptions and an unpredicted rapid cooling 1800 sec into entry is credited to inaccurate assumptions of structural heat dissipative properties or flow conditions in that time phase of the flight; additional discrepancies in descriptions of heating of the upper fuselage are attributed to a lack of knowledge of the complex flow patterns existing over that area of the Shuttle body.

Ko, W. L.↗
NASA NTRS1980

In-flight boundary-layer measurements on a hollow cylinder at a Mach number of 3.0

Skin temperatures, shear forces, surface static pressures, boundary layer pitot pressures, and boundary layer total temperatures were measured on the external surface of a hollow cylinder that was 3.04 meters long and 0.437 meter in diameter and was mounted beneath the fuselage of the YF-12A airplane. The data were obtained at a nominal free stream Mach number of 3.0 (a local Mach number of 2.9) and at wall to recovery temperature ratios of 0.66 to 0.91. The local Reynolds number had a nominal value of 4,300,000 per meter. Heat transfer coefficients and skin friction coefficients were derived from skin temperature time histories and shear force measurements, respectively. In addition, boundary layer velocity profiles were derived from pitot pressure measurements, and a Reynolds analogy factor was obtained from the heat transfer and skin friction measurements. The measured data are compared with several boundary layer prediction methods.

Quinn, R. D.↗
NASA NTRS1978

In-flight Compressible Turbulent Boundary Layer Measurements on a Hollow Cylinder at a Mach Number of 3.0

Skin temperatures, shearing forces, surface static pressures, and boundary layer pitot pressures and total temperatures were measured on a hollow cylinder 3.04 meters long and 0.437 meter in diameter mounted beneath the fuselage of the YF-12A airplane. The data were obtained at a nominal free stream Mach number of 3.0 and at wall-to-recovery temperature ratios of 0.66 to 0.91. The free stream Reynolds number had a minimal value of 4.2 million per meter. Heat transfer coefficients and skin friction coefficients were derived from skin temperature time histories and shear force measurements, respectively. Boundary layer velocity profiles were derived from pitot pressure measurements, and a Reynolds analogy factor of 1.11 was obtained from the measured heat transfer and skin friction data. The skin friction coefficients predicted by the theory of van Driest were in excellent agreement with the measurements. Theoretical heat transfer coefficients, in the form of Stanton numbers calculated by using a modified Reynolds analogy between skin friction and heat transfer, were compared with measured values. The measured velocity profiles were compared to Coles' incompressible law-of-the-wall profile.

Quinn, R. D.↗