Design and optimization of integrated thermal management system for high-speed aircraft

doi:10.11949/0438-1157.20191197

Abstract

Abstract:

Under the dual effects of aerodynamic heating and high-power electronic equipment heating, the advanced high-speed aircraft heat sink and energy demand are exponentially rising, which seriously restricts the function and performance of the aircraft. In order to improve system cooling and power supply performance and reduce engine performance loss, the optimization design of integrated thermal management system is studied. In this paper, based on the large heat load high-speed aircraft with Mach number (Ma)1—4.4, the optimization design for the three integrated thermal management systems is carried out to optimal match the fuel heat sink, the outer duct convection heat sink, the ram air and flight missions. The equivalent quality method was used to analysis, which equalizes the mass, energy consumption and gas source consumption to the fuel weight penalty, and defined as the objective function. The results reveal that when the Mach number is lower than 2, the outer duct air heat sink mode is more economical. However, with the increasing of the flight speed, the refrigeration cycle pressure ratio significantly increased, which cause the fuel weight penalty increased sharply. When the Mach number is 2—4.5, the fuel heat sink mode is more suitable. Its fuel weight penalty is mainly due to the increasing of flight speed. Compared with engine bleed air, rim air is more suitable for higher Mach number. Thus, for the cruise Mach number below 2, the integrated thermal management system equipped with “external duct bleed air heat sink and engine bleed air” earns less engine performance loss. For the cruise Mach number of 2—4.5, the integrated thermal management system equipped with “fuel heat sink and switchable engine bleed air/rim air” earns better engine performance.

Key words: integrated thermal management system, high-speed aircraft, dynamic simulation, thermodynamics, optimization design

摘要：

先进的高速飞行器面临着气动加热与大功率电子设备发热的双重热负荷，使得机载热沉与能量需求呈指数上升趋势，进而导致发动机性能下降、耗油量增加，严重制约着飞行器的功能和性能提升。机载热管理系统的优化设计，旨在提升系统制冷和供电性能的同时减小发动机性能损失。以Mach数Ma=1~4.4的大热负载高速飞行器为背景，针对三种机载综合热管理系统，开展适应飞行任务的系统优化设计，实现燃油热沉、外涵道引气热沉、冲压空气引气、发动机引气与飞行任务的最优匹配。研究过程采用等效质量方法，将各系统质量、能耗、气源消耗等成本统一等效为燃油代偿损失，并作为目标函数，对多种工况进行优化设计。研究结果表明：在Ma≤2时，采用外涵道空气热沉模式更为合适，但随飞行速度的进一步提高，其制冷循环压比显著上升制冷效率降低，燃油代偿损失急剧上升；基于燃油热沉的综合热管理模式更适用于Ma=2~4.5的飞行任务，其制冷循环功耗和能耗在各飞行工况下性能表现较为稳定，燃油代偿损失仅因飞行速度增大而增大；与发动机引气相比，冲压空气引气更适合Mach数较高的飞行任务规划。因此，对于巡航Ma≤2的飞行器，搭载“外涵道引气热沉+发动机引气”的机载综合热管理系统，发动机性能损失更低；对于巡航Ma=2~4.5的飞行器，搭载“燃油热沉+可切换发动机引气/冲压空气引气”的机载综合热管理系统，发动机性能最优。

关键词: 综合热管理系统, 高速飞行器, 动态仿真, 热力学, 优化设计

CLC Number:

V 245.3

Rong A, Liping PANG, Dongsheng YANG, Bin QI. Design and optimization of integrated thermal management system for high-speed aircraft[J]. CIESC Journal, 2020, 71(S1): 315-321.

阿嵘, 庞丽萍, 杨东升, 齐玢. 高速飞行器机载综合热管理系统设计与优化[J]. 化工学报, 2020, 71(S1): 315-321.

Figures/Tables 11

Fig.1 Integrated thermal management system diagram

Table 1 Architecture and Mach number range

1-2-3-4-5-6-1

7(IB)-8-9

外涵道空气

中压级

[1,3)

1-2'-3-4-5-6-1

7(IB)-8-9

高温Pao

中压级

[1,3)

1-2'-3-4-5-6-1

7(RB)-8-9

高温Pao

冲压空气

[1,4.4)

Table 2 Sources and calculations of fuel weight penalty

代偿损失来源	1	2	3	计算方法
进气道阻力	●	○	○	$m b y p a s s = m ˙ a v - v o u t K g e x p C e τ 0 g K - 1$
换热器质量	●	●	●	$m F = M H X e x p C e τ 0 g K - 1$
压缩机质量	●	●	●	$m F = M C e x p C e τ 0 g K - 1$
制冷涡轮质量	●	●	●	$m F = M C T e x p C e τ 0 g K - 1$
发电涡轮质量	●	●	●	$m F = M P T e x p C e τ 0 g K - 1$
冲压空气引气	○	○	●	$m r i m = m ˙ e v K g e x p C e τ 0 g K - 1$
中压级引气	●	●	○	$m E = m ˙ e H * π 0 k - 1 k - 1 H u ε c π c k - 1 k - 1 + m ˙ e v C e K (e C e τ 0 g K - 1) C e g$
燃油消耗	?	?	?	$m f u e l = f m ˙ e K C e g (e C e τ 0 g K - 1)$

Table 2 Sources and calculations of fuel weight penalty

代偿损失来源	1	2	3	计算方法
进气道阻力	●	○	○	$m b y p a s s = m ˙ a v - v o u t K g e x p C e τ 0 g K - 1$
换热器质量	●	●	●	$m F = M H X e x p C e τ 0 g K - 1$
压缩机质量	●	●	●	$m F = M C e x p C e τ 0 g K - 1$
制冷涡轮质量	●	●	●	$m F = M C T e x p C e τ 0 g K - 1$
发电涡轮质量	●	●	●	$m F = M P T e x p C e τ 0 g K - 1$
冲压空气引气	○	○	●	$m r i m = m ˙ e v K g e x p C e τ 0 g K - 1$
中压级引气	●	●	○	$m E = m ˙ e H * π 0 k - 1 k - 1 H u ε c π c k - 1 k - 1 + m ˙ e v C e K (e C e τ 0 g K - 1) C e g$
燃油消耗	?	?	?	$m f u e l = f m ˙ e K C e g (e C e τ 0 g K - 1)$

Table 3 Initial and constraint conditions of optimization

系统		$T 1 / K$	$m ˙ H / k g / s$	$Δ P m a x, c / k P a$	$Δ P m a x, h / k P a$
系统1	x₁	950	10	—	—
	LB	$T z + 50$	0.1	0	0
	UB	1000	20	0.9P_z-P₀	5
系统 2和3	x₁	700	3	—	—
	LB	473	0.1	—	—
	UB	1000	8	—	—

Table 3 Initial and constraint conditions of optimization

系统		$T 1 / K$	$m ˙ H / k g / s$	$Δ P m a x, c / k P a$	$Δ P m a x, h / k P a$
系统1	x₁	950	10	—	—
	LB	$T z + 50$	0.1	0	0
	UB	1000	20	0.9P_z-P₀	5
系统 2和3	x₁	700	3	—	—
	LB	473	0.1	—	—
	UB	1000	8	—	—

Fig.2 Flow of integrated thermal management system optimization

Table 4 Related parameters of optimization

参数	数值
巡航时长/s	4200
高温Pao温度/℃	150
设备最高许用温度/℃	70
闭式循环涡轮出口压力/kPa	100
升阻比	4.62
单位推力燃油消耗量TSFC/h^-1	3600(C_t_,1+ C_t_,2Ma)
常数C_t_,1/s^-1	1.1×10^-4
常数C_t_,2/s^-1	6.8×10^-5
进气道阻力损失系数	0.9
涡轮和压缩机的绝热效率	0.75
机械效率	0.98
发动机涡轮前最高温度/K	1400
发动机压气机末级最高温度/K	800

Fig.3 Curves of system 1 performance changing with Mach number

Fig.4 Results of system 1 minimum fuel weight penalty

Fig.5 Curves of system 2 performance changing with Mach number

Fig.6 Results of system 2 minimum fuel weight penalty

Fig.7 Results of system 3 minimum fuel weight penalty

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