PAT(A) - 1102. Invert a Binary Tree (25)

本文介绍了一个二叉树翻转问题的解决方法,包括如何构建二叉树、找到根节点,并实现层序与中序遍历。通过调整遍历顺序来达到翻转效果。


1102. Invert a Binary Tree (25)

时间限制
400 ms
内存限制
65536 kB
代码长度限制
16000 B
判题程序
Standard
作者
CHEN, Yue

The following is from Max Howell @twitter:

Google: 90% of our engineers use the software you wrote (Homebrew), but you can't invert a binary tree on a whiteboard so fuck off.

Now it's your turn to prove that YOU CAN invert a binary tree!

Input Specification:

Each input file contains one test case. For each case, the first line gives a positive integer N (<=10) which is the total number of nodes in the tree -- and hence the nodes are numbered from 0 to N-1. Then N lines follow, each corresponds to a node from 0 to N-1, and gives the indices of the left and right children of the node. If the child does not exist, a "-" will be put at the position. Any pair of children are separated by a space.

Output Specification:

For each test case, print in the first line the level-order, and then in the second line the in-order traversal sequences of the inverted tree. There must be exactly one space between any adjacent numbers, and no extra space at the end of the line.

Sample Input:
8
1 -
- -
0 -
2 7
- -
- -
5 -
4 6
Sample Output:
3 7 2 6 4 0 5 1
6 5 7 4 3 2 0 1

思路分析:输入结点建立一棵二叉树,关键先找到根节点,然后进行遍历。输出第一行是层序遍历,第二行是中序遍历。然而。。。输出结果怎么和样例不太一样呢。原来还有一个单词是'Invert",意思是翻转。不用真正的做翻转,只要遍历的时候调整一下顺序就OK了。

#include <cstdio>
#include <queue>

#define MAX 10

using namespace std;

typedef struct {
    int index;
    int lchild;
    int rchild;
} Node;

Node node[MAX];
int count = 0;

void dfsInOrder( Node curNode ) {
    if( curNode.rchild != -1 ) {
        dfsInOrder( node[curNode.rchild] );
    }
    if( count == 0 ) {
       printf( "%d", curNode.index );
       count = 1;
    }
    else {
       printf( " %d", curNode.index );
    }
    if( curNode.lchild != -1 ) {
        dfsInOrder( node[curNode.lchild] );
    }

}

int main() {
    //freopen( "123.txt", "r", stdin );
    int isRoot[MAX] = { 0 };
    int n;
    scanf( "%d", &n );
    char ch = getchar();

    for( int i = 0; i < n; i++ ) {
        node[i].index = i;
        char left, right;
        scanf( "%c %c", &left, &right );
        char ch = getchar();

        if( left == '-' ) node[i].lchild = -1;
        else { node[i].lchild = left - '0'; isRoot[left - '0'] = 1; }

        if( right == '-' ) node[i].rchild = -1;
        else { node[i].rchild = right - '0'; isRoot[right - '0'] = 1; }
    }

    //for( int i = 0; i < n; i++ ) {
    //    printf( "%d %d %d\n", node[i].index, node[i].lchild, node[i].rchild );
    //}
    int root = 0;
    for( int i = 0; i < n; i++ ) {
        if( isRoot[i] == 0 ) {
           root = i;
           break;
        }
    }

    //printf( "root: %d\n", root );

    queue<Node> q;
    q.push( node[root] );

    while( !q.empty() ) {
        Node curNode = q.front();
        q.pop();
        if( count == 0 ) {
            printf( "%d", curNode.index );
            count = 1;
        }
        else
            printf( " %d", curNode.index );
        if( curNode.rchild != -1 ) q.push( node[curNode.rchild] );
        if( curNode.lchild != -1 ) q.push( node[curNode.lchild] );
    }
    printf( "\n" );

    count = 0;
    dfsInOrder( node[root] );

    return 0;
}


XFOIL Version 6.99 Calculated polar for: NACA 22012 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.000 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -5.000 -0.4931 0.00000 -0.00117 -0.0003 0.0000 0.0000 -4.000 -0.3724 0.00000 -0.00115 -0.0014 0.0000 0.0000 -3.000 -0.2517 0.00000 -0.00113 -0.0027 0.0000 0.0000 -2.000 -0.1308 0.00000 -0.00111 -0.0040 0.0000 0.0000 -1.000 -0.0099 0.00000 -0.00110 -0.0053 0.0000 0.0000 0.000 0.1110 0.00000 -0.00108 -0.0067 0.0000 0.0000 1.000 0.2318 0.00000 -0.00107 -0.0081 0.0000 0.0000 2.000 0.3526 0.00000 -0.00107 -0.0095 0.0000 0.0000 3.000 0.4733 0.00000 -0.00106 -0.0110 0.0000 0.0000 4.000 0.5938 0.00000 -0.00106 -0.0125 0.0000 0.0000 5.000 0.7142 0.00000 -0.00107 -0.0141 0.0000 0.0000 6.000 0.8343 0.00000 -0.00108 -0.0157 0.0000 0.0000 7.000 0.9542 0.00000 -0.00109 -0.0173 0.0000 0.0000 8.000 1.0738 0.00000 -0.00110 -0.0189 0.0000 0.0000 9.000 1.1930 0.00000 -0.00112 -0.0206 0.0000 0.0000 10.000 1.3119 0.00000 -0.00114 -0.0223 0.0000 0.0000 这是工业软件XFOIL运行得到的结果,你的输出结果尚有差距(venv) PS C:\Users\caoao\Desktop\airfoil_analysis> python aerodynamics_analysis.py ============================================================ NACA 22012 Aerodynamic Analysis Program ============================================================ ============================================================ Analyzing NACA 22012 at α=5.0° with 150 panels ============================================================ 翼型图已保存为 naca22012_airfoil.png Results for α=5.0°: Kutta-Joukowski C_L: -0.020026 Pressure Integration C_L: -0.017880 Thin Airfoil Theory C_L: 0.776106 Zero-Lift Angle (Thin Airfoil): -2.0772° 压力分布图已保存为 pressure_distribution_alpha_5.0.png ============================================================ Analyzing Lift Curve with 150 panels ============================================================ α=-5°: C_L_KJ=-0.0238, C_L_P=0.0367, C_L_thin=-0.3205 α=-4°: C_L_KJ=-0.0234, C_L_P=0.0314, C_L_thin=-0.2109 α=-3°: C_L_KJ=-0.0231, C_L_P=0.0261, C_L_thin=-0.1012 α=-2°: C_L_KJ=-0.0227, C_L_P=0.0207, C_L_thin=0.0085 α=-1°: C_L_KJ=-0.0224, C_L_P=0.0153, C_L_thin=0.1181 α=0°: C_L_KJ=-0.0220, C_L_P=0.0098, C_L_thin=0.2278 α=1°: C_L_KJ=-0.0216, C_L_P=0.0043, C_L_thin=0.3375 α=2°: C_L_KJ=-0.0212, C_L_P=-0.0013, C_L_thin=0.4471 α=3°: C_L_KJ=-0.0208, C_L_P=-0.0068, C_L_thin=0.5568 α=4°: C_L_KJ=-0.0204, C_L_P=-0.0123, C_L_thin=0.6664 α=5°: C_L_KJ=-0.0200, C_L_P=-0.0179, C_L_thin=0.7761 α=6°: C_L_KJ=-0.0196, C_L_P=-0.0234, C_L_thin=0.8858 α=7°: C_L_KJ=-0.0192, C_L_P=-0.0289, C_L_thin=0.9954 α=8°: C_L_KJ=-0.0188, C_L_P=-0.0343, C_L_thin=1.1051 α=9°: C_L_KJ=-0.0183, C_L_P=-0.0397, C_L_thin=1.2148 α=10°: C_L_KJ=-0.0179, C_L_P=-0.0450, C_L_thin=1.3244 升力系数曲线图已保存为 lift_coefficient_curve.png Lift Curve Slope Analysis: Slope (per degree): 0.000391 Slope (per radian): 0.022407 Theoretical (): 6.283185 Analysis complete! Results saved as PNG files.
最新发布
12-09
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